cresis uav critical design review: the meridianred design white design blue design p-3 orion twin...

CReSIS UAV Critical Design Review: The Meridian

William Donovan

University of Kansas 2335 Irving Hill Road

Lawrence, KS 66045-7612 http://cresis.ku.edu

Technical Report CReSIS TR 123

June 25, 2007

This work was supported by a grant from the

National Science Foundation (#ANT-0424589).

i

Executive Summary This report briefly describes the development of the three preliminary configuration

designs proposed for the Meridian UAV. This report details the selection of the

primary configuration and further, more detailed, analysis including Class II weight

and Balance, Class II Stability and Control, Performance Analyses, Systems Design,

Class II landing gear, structural arrangement, a manufacturing breakdown and a cost

analysis.

The design mission for this aircraft is to takeoff from a snow or ice runway, fly to a

designated area, then use low frequency radar to perform measurements of ice sheets

in Greenland and Antarctica. Three designs were developed:

• A Monoplane with Structurally Integrated Antennas

• A Monoplane with Antennas Hanging from the Wing

• A Biplane with Antennas Structurally Integrated Into the Lower Wing

The monoplane with antennas hanging from the wing was selected as the primary

configuration for further development. This report describes the Class II design and

analysis of that vehicle.

ii

Acknowledgments This material is based upon work supported by the National Science Foundation

under Grant No. AST-0424589. Any opinions, findings, and conclusions or

recommendations expressed in this material are those of the author(s) and do not

necessarily reflect the views of the National Science Foundation.

iii

Table of Contents Executive Summary..................................................................................................... i Acknowledgments ....................................................................................................... ii Table of Contents ....................................................................................................... iii List of Figures.............................................................................................................. v List of Tables .............................................................................................................. vi Nomenclature ............................................................................................................ vii Abbreviations ........................................................................................................... viii 1 Summary of Preliminary Designs...................................................................... 9 2 Configuration Selection and Requirement Changes ..................................... 11

2.1.1 Engine Selection – Turboprop Variant ............................................... 12 3 Class II Design................................................................................................... 15

3.1 Class II Weight and Balance....................................................................... 15 3.1.1 The Aircraft V-n Diagram .................................................................. 15

3.2 Component Weight Estimations ................................................................. 18 3.3 Class II Stability and Control...................................................................... 22

3.3.1 Trim Diagrams .................................................................................... 22 3.3.2 Open Loop Dynamics ......................................................................... 31 3.3.3 Actuator Size and Rate Requirements ................................................ 37

3.4 Class II Aerodynamics................................................................................ 38 3.5 Propulsion ................................................................................................... 46 3.6 Performance Analysis ................................................................................. 47

3.6.1 Stall Speed .......................................................................................... 47 3.6.2 Takeoff Distance................................................................................. 47 3.6.3 Climb................................................................................................... 48 3.6.4 Cruise Performance............................................................................. 48 3.6.5 Landing Distance ................................................................................ 50

3.7 Systems ....................................................................................................... 51 3.7.1 Flight Control System......................................................................... 52 3.7.2 Electrical System ................................................................................ 54 3.7.3 Communications/Telemetry System................................................... 58 3.7.4 Fuel System......................................................................................... 59 3.7.5 Anti-Icing System ............................................................................... 60

3.8 Class II Landing Gear ................................................................................. 61 3.8.1 Tire Selection ...................................................................................... 62 3.8.2 Strut Sizing.......................................................................................... 63 3.8.3 Landing Gear Integration.................................................................... 64

3.9 Structural Arrangement............................................................................... 66 3.9.1 Wing Structure .................................................................................... 67 3.9.2 Fuselage Structural Layout ................................................................. 70

3.10 Manufacturing Breakdown ......................................................................... 74 3.11 Cost Analysis .............................................................................................. 75

iv

3.11.1 Research, Development, Test, and Evaluation Costs ......................... 77 3.11.2 Acquisition Cost.................................................................................. 77 3.11.3 Cost Estimate Summary...................................................................... 78 3.11.4 Cost Estimate Justification.................................................................. 81

4 Conclusions........................................................................................................ 84 5 References.......................................................................................................... 85

v

List of Figures Figure 1.1: Comparison of Fuel Usage for 3 Fine Scale Missions ............................. 10 Figure 1.2: Combined Takeoff Weight Regression Chart .......................................... 10 Figure 2.1 - Vivaldi Antenna ...................................................................................... 11 Figure 2.2: CAD Model of Innodyn 165TE................................................................ 13 Figure 3.1 - The Meridian UAV ................................................................................. 14 Figure 3.2 - V-n Diagram for the Meridian ................................................................ 17 Figure 3.3 - Center of Gravity Excursion for the Meridian ........................................ 20 Figure 3.4 – Component C.G. Locations .................................................................... 21 Figure 3.5 - Trim Diagram - Cruise ............................................................................ 23 Figure 3.6 - Trim Diagram - Takeoff, Gear Down ..................................................... 24 Figure 3.7 - Trim Diagram - Takeoff, Gear Up .......................................................... 25 Figure 3.8 - Trim Diagram – Landing Heavy, Gear Down ........................................ 26 Figure 3.9 - Trim Diagram - Landing Heavy, Gear Up .............................................. 27 Figure 3.10 - Trim Diagram - Landing Light, Gear Down......................................... 28 Figure 3.11 - Trim Diagram - Landing Light, Gear Up.............................................. 29 Figure 3.12 - Trim Diagram - OEI.............................................................................. 30 Figure 3.13 - Drag Polars for the Meridian without Antennas ................................... 41 Figure 3.14 - Lift-to-Drag for the Meridian without Antennas .................................. 42 Figure 3.15 - Drag Polars for the Meridian with Antennas ........................................ 43 Figure 3.16 - Lift-to-Drag for the Meridian with Antennas........................................ 44 Figure 3.17 - Relationship of Parasite Area and Wetted Area for Various Single Engine Aircraft [6]...................................................................................................... 46 Figure 3.18 - Cloud Cap Tech. Piccolo II Autopilot [36]........................................... 52 Figure 3.19 - Piccolo II Architecture [36] .................................................................. 53 Figure 3.20 - Piccolo Ground Station and Pilot Controller (Operator Interface Not Shown) [36] ................................................................................................................ 53 Figure 3.21 - Electrical Load Profile for the Meridian UAV ..................................... 56 Figure 3.22 - Electrical System Layout ...................................................................... 57 Figure 3.23 - Fuselage Systems Layout...................................................................... 58 Figure 3.24 - Fuel Tank Integration............................................................................ 60 Figure 3.25 - Lancair Legacy Landing Gear Strut [37] .............................................. 64 Figure 3.26 - Lancair Legacy Landing Gear Installation [38] .................................... 65 Figure 3.27 - Matco Tailwheel Assembly [39]........................................................... 65 Figure 3.28 - Wing Structural Layout......................................................................... 69 Figure 3.29 - Fuselage Structural Layout ................................................................... 71 Figure 3.30 - Wing-Fuselage Attachment................................................................... 72 Figure 3.31 - Standard 20 Foot Shipping Container Door [9] .................................... 73 Figure 3.32 - Typical Engine Mount for the Innodyn 165TE..................................... 74 Figure 3.33 - Manufacturing Breakdown.................................................................... 75 Figure 3.34 - Cost Breakdown by Overall Category .................................................. 80 Figure 3.35 - UAV Cost in Terms of Payload Weight ............................................... 82

vi

Figure 3.36 - UAV Cost Based on System Cost Versus Payload Weight .................. 83

List of Tables Table 1.1: Summary of Preliminary Design Concepts ................................................. 9 Table 3.1 - V-n Diagram Parameters .......................................................................... 16 Table 3.2 - Design Speeds and Load Factors for the Meridian .................................. 16 Table 3.3 - Class II Weight and Balance for the Meridian ......................................... 19 Table 3.4 - Weight and Balance Summary for the Meridian...................................... 19 Table 3.5 - Meridian Flight Conditions ...................................................................... 22 Table 3.6 - Dynamic Analysis Flight Conditions ....................................................... 31 Table 3.7 - Stability Deriviatives for the Meridian..................................................... 33 Table 3.8 - Control Derivatives for the Meridian ....................................................... 34 Table 3.9 - Longitudinal Transfer Functions for Cruise ............................................. 35 Table 3.10 - Lateral Transfer Functions for Cruise .................................................... 35 Table 3.11 - Directional Transfer Functions for Cruise.............................................. 36 Table 3.12 - Meridian Dynamic Stability Parameters ................................................ 36 Table 3.13 - Roll Control Requirements - Time to Achieve Bank Angle (Seconds) . 37 Table 3.14 - Roll Control Results ............................................................................... 37 Table 3.15 Wing Geometry for Drag Calculations..................................................... 38 Table 3.16 - V-Tail Geometry for Drag Calculations................................................. 38 Table 3.17 - Fuselage Geometry for Drag Calculations ............................................. 39 Table 3.18 - Flap and Landing Gear Geometry for Landing Gear Calculations ........ 39 Table 3.19 - Drag Analysis Results ............................................................................ 40 Table 3.20 - Resultant Oswald's Efficiency and Parasite Area for the Meridian ....... 45 Table 3.21 - Stall Speed Summary ............................................................................. 47 Table 3.22 - Landing Gear Strut Sizing ...................................................................... 64 Table 3.23 - Engineering and Manufacturing Rate Estimation .................................. 76 Table 3.24 - RDT&E and Acquisition Cost Summary ............................................... 79 Table 3.25 - Cost Breakdown by Overall Category.................................................... 79 Table 3.26 - Cost Breakdown by RDT&E and Production Categories ...................... 80 Table 3.27 - Current UAV Procurement Cost [40]..................................................... 81

vii

Nomenclature Symbol Description Units

AR Aspect Ratio ~ b Wing Span ft, in c Wing chord ft, in CD Drag Coefficient ~ CD0 Zero-Lift Drag Coefficient ~ CL Lift Coefficient ~ CLα Lift-Curve Slope Rad-1

cp Specific Fuel Consumption Lbs/hp-hr D Drag Lbs Dp Propeller Diameter Ft e Oswald’s Efficiency ~ f Equivalent Parasite Area Ft2 L Lift Lbs M Munk’s Span Factor ~ np Number of Propeller Blades ~ P Engine Power hp Pbl Blade Power Loading hp/ft2

R Range Nm S Wing Area Ft2 SWet Wetted Area Ft2

WE Empty Weight Lbs WF Fuel Weight Lbs Wpay Payload Weight Lbs WTO Takeoff Weight Lbs Γ Dihedral Angle Deg α Angle of Attack Deg ε Wing Twist Deg η Wing Station ~ ηp Propeller Efficiency ~ σ Biplane Interference Factor ~

viii

Abbreviations Abbreviation Description

CReSIS Center for Remote Sensing of Ice Sheets FAR Federal Aviation Regulations NSF National Science Foundation UAV Uninhabited Air Vehicle

9

1 Summary of Preliminary Designs Three preliminary designs have previously been presented to satisfy the

requirements for the CReSIS polar research mission:

• A monoplane with flush-mounted antennas

• A monoplane with hanging antennas

• A biplane with flush mounted antennas

The three designs are summarized in Table 1.1. The fuel required to complete 3

fine-scale mission is shown in Figure 1.1 for the three designs as well as the

Lockheed P-3 and the De Havilland Twin Otter. Figure 1.2 shows the aircraft plotted

on the takeoff weight regression chart.

Table 1.1: Summary of Preliminary Design Concepts

Parameter Units Red Design White Design Blue DesignGeometry

Wing Area ft2 49 82 88Wing Span ft 15.33 25.6 17.2Length Overall ft 16 17.5 16.5Height Overall ft 5.5 5.6 5.6

WeightsTakeoff Weight lbs 760 1,270 950Empty Weight lbs 450 720 550Payload Weight lbs 121 121 121Fuel Weight lbs 185 425 270

PerformanceRange nm 1,750 1,750 1,750L/DCr ~ 12.5 8.0 10.0

PowerplantEngine ~ Rotax 912-A Rotax 914-F Rotax 914-FPower hp 81 115 115

10

8,000

1,200

188

82

119

0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Red

Des

ign

Whi

te D

esig

nB

lue

Des

ign

P-3

Orio

nTw

in O

tter

Fuel Required for 3 Fine Missions, gallons

Figure 1.1: Comparison of Fuel Usage for 3 Fine Scale Missions

10

100

1,000

10,000

100 1,000 10,000

Takeoff Weight, lbs

Empt

y W

eigh

t, lb

s

log10(WTO) = A + B*log10(WE) A = -0.0183 B = 1.0930

Dakota

Shadow 200

Shadow 600

I-Gnat

E-Hunter

Predator

Predator B

Figure 1.2: Combined Takeoff Weight Regression Chart

11

2 Configuration Selection and Requirement Changes The three preliminary designs presented were compared based on several criteria

including weight, aerodynamic performance, and antenna integration. The selection

of the configuration for further study was made primarily based on antenna

integration issues. While the Red and Blue designs were more efficient than the

White design aerodynamically, the antenna integration methods used for these

designs would limit the possible bandwidth of the radar, thereby decreasing the total

system performance. The success of the CReSIS project is heavily dependent on the

level of synergy between the radar and aircraft systems. Design decisions must be

made based on high level, systematic concerns.

Figure 2.1 - Vivaldi Antenna

L = 0.5 m H = 0.5 m T = 0.125 m

12

The White design was therefore selected as the configuration for further

development. However, the antenna type must be changed to accommodate higher

operating bandwidths. The primary antenna will be a Vivaldi antenna as shown in

Figure 2.1. These are essentially flat plates that hang from the wing of the aircraft.

While these have been selected as the current antenna for the radar system, it is clear

that the antenna type may change again. Therefore, the White design was modified

slightly such that eight mounting points will be integrated into the wing structure

from which several types of antennas could be mounted. The aircraft will be

designed for the Vivaldi antennas, but if another type of antenna of equal or smaller

size proves to be more effective, then it may be mounted on the wing without any

structural modifications. Essentially, the new design philosophy is that the aircraft is

relatively insensitive to the type and size of antennas used within reason. This

solution will yield a highly adaptable, high performance vehicle capable of multiple

missions.

2.1.1 Engine Selection – Turboprop Variant

The engines selections shown for the three preliminary designs were driven by

power and specific fuel consumption requirements. The reliability of these engines in

a cold-weather environment is questionable. Also, from a logistics standpoint, the

Rotax engines are suboptimal as the primary fuel used in Antarctica is Jet-A, not

aviation gas. For these reasons, the Innodyn 165TE (Figure 2.2) has been selected for

further investigation. The Innodyn is a fairly new, small turbopropeller engine that

13

has yet to be fully tested. Therefore, the specific fuel consumption of the engine is

somewhat unknown.

Figure 2.2: CAD Model of Innodyn 165TE

The current specific fuel consumption estimate for the Innodyn engine is 0.70

lbs/hp-hr, which is much higher than the Rotax 912 and 914 value of 0.56 lbs/hp-hr.

Nonetheless, the reliability and maintainability issues make this engine very

appealing for this mission. The Innodyn has been selected as the primary engine for

the Meridian pending testing that will be performed in the Fall of 2006. The Rotax

914 will be considered as a backup to the Innodyn in the case of poor test results.

14

Figure 2.3 - The Meridian UAV

15

3 Class II Design The purpose of this document is to expand upon the chosen aircraft configuration

(Monoplane with antennas hanging from the wing) through Class II design. The new

Meridian design shown in Figure 2.3 is the result of several design iterations focused

on manufacturability and operational constraints.

3.1 Class II Weight and Balance

The purpose of this section is to describe the Class II weight and balance performed

for the Meridian UAV. This consisted of first calculating and plotting a V-n diagram

to determine the limit and ultimate loading for the Meridian. The results of the V-n

diagram were then used to create weight estimates for each vehicle component.

Finally, a weight and balance analysis is presented to show the aircraft center of

gravity travel.

3.1.1 The Aircraft V-n Diagram

A V-n diagram was constructed for the Meridian UAV to help determine the

maximum load factors and design speeds that will be used for structural sizing. The

V-n diagram was created based on FAR 23 requirements for Normal class aircraft as

there are currently no certification requirements for UAVs. The inputs to the V-n

diagram creation are shown in Table 3.1.

16

Table 3.1 - V-n Diagram Parameters

Parameter Value UnitsAltitude 0 ftWgross 1,125 lbs

S 69.6 ft2

W/S 16.2 psfm.g.c. 2.64 ft

Cla 3.98 rad-1

CLmax (+) 1.3 ~CLmax (-) -0.97 ~

CD @ CLmax (+) 0.085 ~CD @ CLmax (-) 0.064 ~

The V-n diagram for the Meridian is shown in Figure 3.1. The design speeds and

limit load factors are shown in Table 3.2. The positive load factor was set to 3.8

based on FAR 23 requirements [35]. The negative load factor was set to 40 percent

of the positive load factor according to FAR 23 requirements [35].

Table 3.2 - Design Speeds and Load Factors for the Meridian

Parameter Value UnitsVs 61 ktsVC 133 ktsVD 186 ktsVA 118 kts

VS,neg 66 ktsnlimit (+) 3.8 ~nlimit (-) -1.5 ~

17

-3

-2

-1

0

1

2

3

4

0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200

Speed, KEAS

Load

Fac

tor,

n

+VC Gust Line

+VD Gust Line

-VC Gust Line

-VD Gust LineNegative g Limit = -1.5

Positive g Limit = 3.8

VS VA

VC VD

Figure 3.1 - V-n Diagram for the Meridian

18

3.2 Component Weight Estimations

Four methods were used for estimating aircraft component weights: the Cessna,

Torenbeek, General Dynamics, and USAF methods. These methods are integrated

into the AAA software, which was used for the weight estimations [7].

These methods are designed for conventional, inhabited aircraft, which the

Meridian is not. For this reason a certain amount of designer intuition was employed

to select the most applicable methods for each component. For example, the Cessna

method produced a wing weight of approximately 300 lbs, while the USAF and

Torenbeek methods resulted in weights of approximately 100 lbs. The latter results

were deemed to be reasonable, therefore the Cessna method was not used for the

wing weight estimation. Table 3.3 shows the component weights as well as the

methods used. The data shown in Table 3.3 are the result of several iterations.

19

Table 3.3 - Class II Weight and Balance for the Meridian

Method Class II Weight XCG YCG ZCG

lbsStructure

Wing USAF, Torenbeek 100.0 101.6 0.0 40.0Empennage Cessna, USAF 22.5 224.0 0.0 50.0Fuselage Cessna, USAF 59.3 121.2 0.0 50.0Landing Gear

Main Gear Cessna 51.5 96.0 0.0 38.0Tail Wheel Cessna 9.0 220.0 0.0 48.0Main Gear - Retracted Cessna 51.5 101.0 0.0 42.0Tail Wheel - Retracted Cessna 9.0 225.0 0.0 48.0

PropulsionPropeller Torenbeek/GD 34.8 50 0 50Engine Manufacturer 188.0 64.0 -0.8 48.0Fuel System USAF, Torenbeek 31.7 102.0 0.0 40.0Engine Systems Torenbeek/GD 43.6 70.0 0.0 0.0

Fixed EquipmentFlight Control System Cessna, Torenbeek 22.4 120.0 0.0 50.0Avionics/Electronics Class I/Manufacturer 11.2 120.0 0.0 50.0Electrical System Cessna, Torenbeek 24.4 118.0 0.0 50.0Icing System USAF, Torenbeek 15.6 120.0 0.0 50.0Paint Torenbeek 3.6 120.0 0.0 50.0

Fuel and PayloadMission Fuel 239.9 107.0 0.0 50.0Fuel Reserves 54.0 107.0 0.0 50.0Trapped Fuel and Oil 5.4 107.0 0.0 50.0Payload 165.0 104.0 0.0 50.0

TotalsStructure - Gear Extended 242.3 120.97 0.00 43.25Structure - Gear Retracted 242.3 122.22 0.00 44.10Powerplant 298.1 67.28 -0.50 40.36Fixed Equipment 77.2 119.37 0.00 50.00Empty Weight 617.6 94.86 -0.13 42.34Useful Load 464.3 105.93 0.00 50.00Total - Gear Extended 1081.9 97.88 -0.13 44.56Total - Gear Retracted 1081.9 98.06 -0.13 44.69

The c.g. locations of each component are shown in Figure 3.3. The c.g. travel due

to fuel and payload loading is shown in Table 3.4 and Figure 3.2.

Table 3.4 - Weight and Balance Summary for the Meridian

Parameter Inches % mgcMost Forward c.g. 94.86 0.18

Most Aft c.g. 99.89 0.34Total Excursion 5.03 0.16Fuel Excursion 2.76 0.09

20

600.0

700.0

800.0

900.0

1000.0

1100.0

1200.0

94 95 96 97 98 99 100

Fuselage Station, in

Wei

ght,

lbs

0.16 0.17 0.18 0.19 0.2 0.21 0.22 0.23 0.24 0.25 0.26 0.27 0.28 0.29 0.3 0.31 0.32 0.33 0.34 0.35

Wing Chord, %mgc

+/- Payload

+ Fuel

- Fuel

Trapped Fuel and Oil

Retract Gear

Extend Gear

WTO

WOE

WE

Figure 3.2 - Center of Gravity Excursion for the Meridian

21

Figure 3.3 – Component C.G. Locations

22

3.3 Class II Stability and Control

The purpose of this section is to describe the Class II stability and control analyses

performed for the Meridian UAV. These include:

• Trim Diagram (Power on and power off)

• Roll Performance

• Crosswind Control During Final Approach and While on the Runway

• Open Loop Dynamic Handling

• Actuator Size and Rate Requirements

3.3.1 Trim Diagrams

Trim diagrams were created for the flight conditions listed in Table 3.5. The trim

diagrams shown in Figure 3.4 through Figure 3.11 were created using the AAA

software [7]. The V-tail incidence was adjusted to ivee = -3.5 deg so that the aircraft

could be trimmed at the stall speed with the most forward center of gravity.

Table 3.5 - Meridian Flight Conditions

Flight Condition Altitude Speed Weight Flaps Gearft kts lbs deg ~

Clean 5,000 120 963 0 UpTakeoff Gear Down 0 60 1,082 0 Down

Takeoff Gear Up 0 60 1,082 0 UpLanding Heavy, Gear Down 0 65 1,082 30 Down

Landing Heavy, Gear Up 0 65 1,082 30 UpLanding Light, Gear Down 0 65 843 30 Down

Landing Light, Gear Up 0 65 843 30 UpOEI 5,000 80 963 0 Up

23

Figure 3.4 - Trim Diagram - Cruise

24

Figure 3.5 - Trim Diagram - Takeoff, Gear Down

25

Figure 3.6 - Trim Diagram - Takeoff, Gear Up

26

Figure 3.7 - Trim Diagram – Landing Heavy, Gear Down

27

Figure 3.8 - Trim Diagram - Landing Heavy, Gear Up

28

Figure 3.9 - Trim Diagram - Landing Light, Gear Down

29

Figure 3.10 - Trim Diagram - Landing Light, Gear Up

30

Figure 3.11 - Trim Diagram - OEI

31

The trim diagrams shown in Figure 3.4 through Figure 3.11 show that the aircraft

can be trimmed throughout the entire flight envelope requiring no more than 20

degrees of control surface deflection.

3.3.2 Open Loop Dynamics

The open loop dynamics were calculated for the Meridian using the AAA program

[7]. The longitudinal and lateral-directional dynamics and flying qualities were

calculated for the takeoff, cruise, and approach flight conditions as specified in Table

3.6. These were compared to the flying quality requirements specified in MIL-F-

8785C [6] and MIL-STD-1797A [6] for a Class I aircraft. While it is not necessary

for a UAV to meet the military specifications, it is common design practice to use the

military flying quality requirements as a basis for the dynamic analysis.

Table 3.6 - Dynamic Analysis Flight Conditions

Parameter UnitsTakeoff Cruise Approach

Altitude ft 0.00 5000.00 0.00ΔT deg F 0.00 -40.00 0.00U1 kts 60.00 120.00 65.00W lbs 1083.4 962.9 843.2α deg 13.28 0.63 6.82

CL1 ~ 1.15 0.30 0.84n g 1.00 1.00 1.00δF deg 0.00 0.00 40.00Xcg in 99.97 99.21 98.25Zcg in 45.66 45.12 44.42εvee deg 0.96 0.41 1.28

ηvee, p. off ~ 1.00 1.00 1.00ηvee ~ 1.96 1.15 1.00

Flight Condition

32

The stability and control derivatives for Meridian are shown in Table 3.7 and Table

3.8, respectively. These values represent the result of several iterations of control

surface sizing.

33

Table 3.7 - Stability Deriviatives for the Meridian


CTx ~ 0.54 0.08 0.05CMT ~ -0.072 -0.011 -0.007CDu ~ 0 0 0CLu ~ 0.010 0.012 0.008CMu ~ 0.002 0.003 0.002CTXu ~ -1.61 -0.23 -0.135027CMTu ~ 0.22 0.03 0.0237792CDα rad-1

0.39 0.10 0.29CLα rad-1

4.01 4.06 4.01CMα rad-1

-0.42 -0.52 -0.64CMTα rad-1

-0.46 -0.31 -0.07CDα rad-1

0 0 0CLα rad-1

0.53 0.55 0.54CMα rad-1

-2.09 -2.17 -2.16CDq rad-1

0 0 0CLq rad-1

3.58 3.82 4.03CMq rad-1

-9.63 -9.85 -9.94CYβ rad-1

-0.42 -0.42 -0.42Clβ rad-1

-0.13 -0.09 -0.12Cnβ rad-1

0.11 0.11 0.11CnTβ rad-1

-0.001 -0.001 -0.001CYβ rad-1

0.0125 0.0003 0.0107Clβ rad-1

-0.0005 0.0000 0.0000Cnβ rad-1

0.0051 0.0001 0.0044CYP rad-1

-0.06 -0.12 -0.09ClP rad-1

-0.46 -0.46 -0.46CnP rad-1

-0.16 -0.04 -0.10CYr rad-1

0.27 0.27 0.27Clr rad-1

0.31 0.10 0.20Cnr rad-1

-0.13 -0.11 -0.12CDivee rad-1

0.01 0.01 0.01CLivee rad-1

0.30 0.30 0.30Cmivee rad-1

-1.17 -1.19 -1.19

Flight Condition

34

Table 3.8 - Control Derivatives for the Meridian


CDδrv rad-10.003 0.004 0.006

CLδrv0 rad-10.14 0.14 0.14

CLδrv rad-10.05 0.14 0.14

CMδrv0 rad-1-0.56 -0.56 -0.56

CMδrv rad-1-0.18 -0.56 -0.56

Chβrv rad-1-0.03 -0.07 -0.03

Chδrv rad-1-0.28 -0.35 -0.29

CYδa rad-10 0 0

Clδa rad-10.12 0.13 0.12

Cnδa rad-1-0.03 -0.01 -0.02

Chαa rad-10.21 0.18 0.21

Chδa rad-10.04 -0.03 0.03

CL0 ~ 0.26 0.26 0.26CL0 ~ 0.26 0.26 0.39CM0 ~ 0.002 -0.003 -0.021Cm0wf ~ -0.05 -0.05 -0.08

Flight Condition

The stability and control derivatives shown in Table 3.7 and Table 3.8 were used to

calculate the open loop transfer functions for the Meridian. This was done with the

AAA software [7]. The transfer functions for the Cruise condition are shown in

Table 3.9 through Table 3.11. The dynamic stability parameters related to the aircraft

flying qualities are shown in Table 3.12. The Meridian met Level I flying quality

requirements for all flight conditions.

35

Table 3.9 - Longitudinal Transfer Functions for Cruise

Table 3.10 - Lateral Transfer Functions for Cruise

36

Table 3.11 - Directional Transfer Functions for Cruise

Table 3.12 - Meridian Dynamic Stability Parameters

Units Takeoff Cruise ApproachLongitudinal

ωsp rad/s 2.9 4.63 2.31ζsp ~ 0.57 0.45 0.59ωp rad/s 0.34 0.25 0.2ζp ~ 0.23 0.1 0.09n/α g/rad 3.2 12.1 5.8CAP 1/gsec2

2.63 1.77 0.92Lateral-Directional

TCSpiral sec -25.6 106.5 -650.7TCroll sec 0.3 0.16 0.27ωD rad/s 2.2 2.9 2.1ζD ~ 0.12 0.1 0.09

Note: Level I flying qualities met for all flight conditions.

Flight ConditionParameter

Roll Control Effectiveness The roll control effectiveness is a vital parameter for aircraft controllability,

especially during approach and landing. The roll control requirements for a Class I

37

aircraft are specified in Table 3.13. The requirement states that the aircraft must be

able to achieve the specified bank angle in the specified amount of time.

Table 3.13 - Roll Control Requirements - Time to Achieve Bank Angle (Seconds)

Cat A Cat B Cat CLevel φt = 60 deg φt = 60 deg φt = 30 deg

I 1.3 1.7 1.3II 1.7 2.5 1.8III 2.6 3.4 2.6

The roll control analysis results are shown in Table 3.14. These values were

calculated using AAA [7]. As can be seen, the Meridian meets Level I flying

qualities for all flight conditions.

Table 3.14 - Roll Control Results

Takeoff Cruise ApproachCat. C B C

φt 30 60 30tr 1.25 1.1 1.2

3.3.3 Actuator Size and Rate Requirements

The actuators were sized using the AAA program to calculate the hinge moment

derivatives. The actuators were then sized for a factor of safety of 2.0 at maximum

deflection. The most critical actuator size was determined to be the flaps, which

required a servo with a maximum torque of at least 100 in-lbs. The servo selected is

the Model 820 manufactured by Moog Components Group (www.polysci.com). This

servo has a peak torque of 150 in-lbs and accepts a PWM signal, which is compatible

with the selected autopilot.

38

3.4 Class II Aerodynamics

The Class II drag analysis was performed for the Meridian with the AAA software

[7]. The geometries used in the drag calculations are shown in Table 3.15 through

Table 3.18. The drag analysis was performed for the takeoff, cruise, approach, and

OEI flight conditions with and without antennas. These flight conditions are

described in Table 3.5.

Table 3.15 Wing Geometry for Drag Calculations

Parameter Units ValueS ft2 69.6

AR ~ 10λ ~ 1

Λc/4 deg 0(t/c)r % 18(t/c)t % 18

LER/c % 1.5L' ~ 2

xlam/c % 10

Clα rad-1 4.3fgap ~ 0.97

ksand 10-3ft 0.00167

Table 3.16 - V-Tail Geometry for Drag Calculations

Parameter Units ValueS ft2 6.1

AR ~ 4λ ~ 0.5

Λc/4 deg 26.3(t/c)r % 12(t/c)t % 12

LER/c % 1.58L' ~ 2

xlam/c % 10Clα rad-1 6.25fgap ~ 0.96

ksand 10-3ft 0.00167

39

Table 3.17 - Fuselage Geometry for Drag Calculations

Parameter Units ValueSb ft2 0.01

Swet ft2 75.1L ft 14.8

Sfrontal ft2 3xlam/L % 0Swet-lam ft2 0

Splf ft2 23.9Df max ft 24ksand 10-3ft 0.00167

Table 3.18 - Flap and Landing Gear Geometry for Landing Gear Calculations

Parameter Units ValueFlaps

Flap Type ~ Plainηi f % 27ηo f % 57

cf/cw % 25Main Gear

Sft ft2 0.25Lstrut ft 2.5

TailwheelCDref ~ 0.5Sref ft2 0.02FD ~ 0

40

Table 3.19 - Drag Analysis Results

Component Takeoff Cruise Approach OEIZero-Lift Drag

Wing 0.0124 0.0107 0.0122 0.0117Vee Tail 0.0021 0.0018 0.0021 0.002Fuselage 0.0018 0.0034 0.0018 0.0041Flap 0 0 0.0188 0Retract 0.0108 0 0.0108 0Fixed Gear 0.0001 0.0001 0.0001 0.0001Trim 0.0018 0.0005 0.0015 0.0004Propeller 0 0 0 0.0267Inlet 0.0001 0.0001 0.0001 0.0001Nozzle 0.0015 0.0015 0.0021 0.0015Power 0.0064 0.0004 0.0013 0Gear Pod 0.0014 0.0012 0.0014 0.0013

Total Zero-Lift 0.0384 0.0197 0.0522 0.0479Drag Due to Lift

Wing 0.0548 0.0034 0.0254 0.0142Vee Tail 0.0009 0 0 0Fuselage 0.0051 0 0.0006 0

Total Due to Lift 0.0608 0.0034 0.026 0.0142Total Drag 0.0992 0.0231 0.0782 0.0621

Antenna Drag (6) 0.0042 0.0037 0.0042 0.0042Total Drag w/ Ant. 0.1034 0.0268 0.0824 0.0663

The drag results are shown in Table 3.19 and Figure 3.12 through Figure 3.15. The

mid-cruise lift-to-drag ratio was found to be 14.5 as indicated on Figure 3.13. This is

slightly less than the value of 16.0 estimated in the Class I design. However, the

performance analysis results show that the range requirement is still met in the

current configuration so no further iteration is necessary.

41

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0 0.05 0.1 0.15 0.2 0.25

Drag Coefficient, CD

Lift

Coe

ffici

ent,

CL

TakeoffCruiseApproachOEI

Reference Data: S = 68.6 ft2

AR = 10.0Takeoff: W = 1,083 lbs e = 0.84 f = 2.84 ft2

Cruise: W = 963 lbs e = 0.83 f = 1.4 ft2

Approach: W = 843 e = 0.75 f = 3.7 ft2

OEI: W = 963 e = 0.81 f = 3.8 ft2

Figure 3.12 - Drag Polars for the Meridian without Antennas

42

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0 2 4 6 8 10 12 14 16 18 20


Lift

Coe

ffici

ent,

CL




Cruise: W = 963 lbs e = 0.83 f = 1.4 ft2

Approach: W = 843 e = 0.75 f = 3.7 ft2

OEI: W = 963 e = 0.81 f = 3.8 ft2

Figure 3.13 - Lift-to-Drag for the Meridian without Antennas

43

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0 0.05 0.1 0.15 0.2 0.25


Lift

Coe

ffici

ent,

CL




Cruise: W = 963 lbs e = 0.83 f = 1.65 ft2

Approach: W = 843 e = 0.75 f = 3.9 ft2

OEI: W = 963 e = 0.81 f = 4.13 ft2

Figure 3.14 - Drag Polars for the Meridian with Antennas

44

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

0 5 10 15 20 25


Lift

Coe

ffici

ent,

CL




Cruise: W = 963 lbs e = 0.83 f = 1.65 ft2

Approach: W = 843 e = 0.75 f = 3.9 ft2

OEI: W = 963 e = 0.81 f = 4.13 ft2

Figure 3.15 - Lift-to-Drag for the Meridian with Antennas

45

To verify the validity of the drag analysis the Oswald’s efficiency and parasite area

were calculated for each flight condition using Equation 3.1 and Equation 3.2

respectively. These values are shown on the drag polar plots in Figure 3.12 through

Figure 3.15 and Table 3.20. The cruise values of the parasite area with and without

the antennas were plotted against known aircraft in Figure 3.16 from [6] for further

verification. As can be seen, Meridian falls somewhere between the lines for overall

skin friction coefficient values of Cf = 0.005 to 0.006 depending on whether the

antennas are installed or not. This is a good indication that the Class II drag results

are reasonable.

eACC

e

L

D2

1

δδ

= Equation 3.1

wD SCf0

= Equation 3.2

Table 3.20 - Resultant Oswald's Efficiency and Parasite Area for the Meridian

No Antennas With AntennasSwet = 240 ft2 Swet = 275 ft2

Flight Condition e f f~ ft2 ft2

Takeoff 0.84 2.84 3.13Cruise 0.82 1.39 1.65Approach 0.75 3.68 3.97OEI 0.81 3.84 4.13

46

Figure 3.16 - Relationship of Parasite Area and Wetted Area for Various Single

Engine Aircraft [6]

3.5 Propulsion

The installed thrust of the Innodyn engine was calculated using the AAA program

[7]. The Innodyn is rated at 165 SHP. The extracted power is estimated at 5 hp based

on the electrical power generation required. The total installed power is 125 hp,

which is above what is required for takeoff and climb performance.

The AAA program was also used to calculate an estimate for the inlet area. This

resulted in an inlet with an area of 0.2 ft2.

47

3.6 Performance Analysis

The purpose of this section is to describe the performance requirements imposed on

this aircraft design and to verify that these requirements have been met by the

Meridian. This includes:

• Stall speed

• Takeoff Distance

• Cruise Performance

• Landing Distance

3.6.1 Stall Speed

The stall speed of the Meridian was calculated for heavy and light flight conditions

with zero and full flaps as shown in Table 3.21. The maximum trimmed lift

coefficients determined in Section 3.3.1 were used for the clean and full flap

configurations. Power effects were ignored for the stall speed calculations.

Table 3.21 - Stall Speed Summary

Parameter Units Light Heavy Light HeavyWeight lbs 843 1,082 843 1,082Flaps deg 0 0 30 30CLmax ~ 1.42 1.42 1.70 1.70Altitude ft 0 0 0 0Vs ft/s 50 57 46 52

Clean Full Flaps

3.6.2 Takeoff Distance

The takeoff distance for the Meridian was calculated for conventional tires as well

as ski operations using the AAA program [7]. This process uses methods found in [6]

48

to calculate the takeoff distance to clear an obstacle of a specified height. The

following assumptions were used in the takeoff distance calculations:

• Ground Friction Coefficient ( 02.0=Gμ for Tires, 15.0=Gμ for skis)

• The obstacle height is 50 ft

• Weight = 1,082 lbs

• Standard sea level conditions

• Drag is based on Class II drag analysis for Takeoff condition:

041.00 =DC

• V3/VTO = 1.3 (Based on FAR 23)

The takeoff analysis resulted in the following takeoff distances:

• Standard Tires on Asphalt: STO = 415 ft

• Skis on Snow: STO = 635 ft

The takeoff distances with and without skis both exceed the required distance of

1,500 ft by a large amount. This is due to the fact that the selected engine has more

power than required by performance matching.

3.6.3 Climb

3.6.4 Cruise Performance

The cruise performance calculations for the Meridian were performed with the

AAA program [7]. This consisted of estimating the range and endurance assuming

constant speed cruise.

49

Range

The range of the Meridian was calculated using the constant speed range equation

found in [6]. The lift-to-drag value was calculated using the mid-cruise Class II drag

polar (Section 3.4). The following assumptions were used for the range calculation:

• Wbegin = 1,083 lbs

• Wfuel = 240 lbs, Wfuel, res = 55

• ηp = 0.80

• cp = 0.90 lbs/hp-hr

The range of the Meridian was determined to be:

• Without Antennas

940 nm (1,735 km) without reserves

1,200 nm (2,220 km) with fuel reserves

• With 6 Antennas

830 nm (1,530 km) without fuel reserves

1,030 nm (1,900 km) with fuel reserves

The range for the Meridian without antennas is acceptably close to the required

range of 1,750 km without antennas and the range with antennas exceeds the required

1,500 km.

Endurance

The endurance of the Meridian was calculated using Equation 3.3. The following

assumptions were made for the endurance calculation:

• Constant speed cruise

50

• Wbegin = 1,083 lbs

• Wfuel = 240 lbs, Wfuel, res = 55 lbs

• ηp = 0.80

• cp = 1.2 lbs/hp-hr (From manufacturer data)

• U1 = 80 kts

• Drag based on mid cruise Class II drag polar

⎥⎥⎦

⎤

⎢⎢⎣

⎡⎟⎟⎠

⎞⎜⎜⎝

⎛

−⎟⎠⎞

⎜⎝⎛

⎟⎠⎞

⎜⎝⎛=

fuelbegin

begin

p

p

WWW

DL

UcE ln

688.155060

1

η Equation 3.3

The loiter speed was set at 80 kts as this is the speed for maximum L/D. The

results from the endurance calculations are:

• Without Antennas

12.7 hours without reserves

16.1 hours with fuel reserves

• With 6 Antennas

11.5 hours without fuel reserves

14.8 hours with fuel reserves

The Meridian exceeds the specified endurance requirements with and without the

antennas and fuel reserves.

3.6.5 Landing Distance

The landing distance for the Meridian was calculated with the AAA program [7].

The landing distance includes the distance from a 50 ft obstacle to the ground and the

51

distance from touchdown to a full stop. The following assumptions were used for the

landing gear calculation:

• Weight = 1,082 lbs

• CLmax = 1.7 (Based on maximum trimmed lift coefficient)

• Drag based on Class II drag polar for the Approach flight condition

• Average ground acceleration = 0.45 g

• Δn = 0.10 (Correction factor due to pilot technique)

The results of the landing distance calculation are:

• Sair = 1,110 ft (Distance in air from obstacle to ground)

• SLG = 430 ft (Ground run distance)

• SL = 1,540 ft (Total distance)

The total landing distance is acceptably close to the required landing distance of

1,500 ft. This distance is based on conventional tires, which was determined to be the

critical requirement as skis actually have a higher coefficient of friction than wheels

on asphalt.

3.7 Systems

The purpose of this section is to describe the systems both on and off the Meridian

that are required for operation. These include:

• Flight Controls

• Electrical System

• Communications

52

• Fuel

• Anti-Icing

3.7.1 Flight Control System

The Meridian will utilize a fly-by-wire control system based around the Piccolo

autopilot, which is produced by Cloud Cap Technologies [36]. The Piccolo requires

dynamic and static pressure inputs and electrical power. The Piccolo interfaces with

the servo actuators using a Pulse Width Modulated (PWM) signal, which is standard

for remote control aircraft. The architecture for the Piccolo is shown in Figure 3.18.

Figure 3.17 - Cloud Cap Tech. Piccolo II Autopilot [36]

53

Figure 3.18 - Piccolo II Architecture [36]

The ground equipment associated with the Piccolo autopilot consists of a ground

station, operator interface (PC), and a pilot control unit (Futaba Controller) as shown

in Figure 3.19.

Figure 3.19 - Piccolo Ground Station and Pilot Controller (Operator Interface Not

Shown) [36]

54

3.7.2 Electrical System

This section describes the electrical system layout for the Meridian including an

electrical load profile for a typical mission. The Meridian will require both 12 and

24VDC power busses. The power system consists of:

• Electrical Generator

• Battery

• Electrical Bus

• Electrical Wiring

The first step in developing the electrical system layout was to generate an

electrical load profile for the Meridian. This was done by listing all necessary

systems required during each phase of a given flight. The results of the load profile

are shown in Figure 3.20. The total load was estimated assuming the radar system is

turned on at takeoff, while the essential load assumes the radar system only requires

power during the on-station flight phase. The most critical flight phases are the

takeoff and landing segments as the landing gear and flap actuators will be operated

in addition to the other systems. The emergency flight phase is representative of an

engine flame-out situation. The battery was sized such that all necessary systems

could remain operating will the aircraft descends and attempts tot restart the engine.

This however, will require the ability to turn some systems off autonomously.

The current configuration of the Innodyn engine is with one 600W, 12 V generator

and a separate starter. The current generator is a standard off-the-shelf automotive

55

alternator, and can be replaced with a larger generator for the Meridian. The

electrical load profile indicates that a 1,000 W generator would be sufficient.

The electrical system layout is shown in Figure 3.21. The wiring is not shown in

Figure 3.21 for clarity. The aileron and flap servo and landing gear actuator wiring

will be located just aft of the aft spar. The antenna wiring will be located just behind

the forward spar. A more detailed view of the systems located in the fuselage are

shown in Figure 3.22.

56

0

100

200

300

400

500

600

700

800

900

1000

Load

ing Start

Taxi

Takeo

ff

Climb

Cruise

Out

On Stat

ion

Cruise

In

Desce

nt

Land

Emergen

cy

Flight Segment

Elec

tric

al P

ower

, W

Total LoadEssential Load

15 Min 5 Min 5 Min 5 Min 20 Min 105 Min 540 Min 105 Min 15 Min 5 Min 10 Min

Figure 3.20 - Electrical Load Profile for the Meridian UAV

57

Figure 3.21 - Electrical System Layout

58

Figure 3.22 - Fuselage Systems Layout

3.7.3 Communications/Telemetry System

The Meridian will utilize dual line-of-site communication links: the piccolo

communications will be used for command and control and a secondary

communications link will be used for vehicle health monitoring/telemetry. For

beyond line-of-site (BLOS) communications, an Iridium satellite communication link

will be utilized. The Piccolo autopilot is configured to transmit and receive data over

an Iridium link. This communications link will be used for low-bandwidth health

monitoring and limited control.

59

3.7.4 Fuel System

The fuel tank integration was difficult for the Meridian due to the removable wing

design. Several options were investigated including a hinged wing joint such that the

wing pivots rearward but is not removed. This would theoretically allow for fuel to

be placed in the outboard wing section, but this type of fuel system would have an

extremely high probability of leaking. Another option is to use quick fuel line

connectors at the wing split. Again, this type of integration poses serious leaking

problems. The design was iterated such that the fuel could be stored inboard of the

wing split. This involved increasing the wing thickness to an 18 percent thick airfoil

and adding a tank in the fuselage. The latter decision required the fuselage height to

grow.

The required fuel volume is 43.7 gallons or 5.84 ft3. Approximately 45 gallons of

fuel fits in the fuel tanks in the inboard wing and fuselage sections as shown in Figure

3.23. Fuel bladders will be utilized for the wing and fuselage tanks. These bladders

are commercially available and include all of the pickups, lines, and baffling as

required. The fuel tank will be split into 3 separate bladders in the wing (1 center,

and two outboard of the inboard rib), and 1 bladder in the fuselage. The center wing

bladder will serve as the fuel collection point.

60

Figure 3.23 - Fuel Tank Integration

3.7.5 Anti-Icing System

A combination of muffed engine exhaust and electrically heated elements will be

utilized for the anti-icing system. The location of the wing is such that the leading

edge of the wing is forward of the firewall. Similar engine installations have been

performed (www.innodyn.com) utilizing NACA inlets to pressurize the engine cowl

volume. This air will be pushed through a valve into the leading edge of the wing.

The temperature of the muffed exhaust air has been measured at 180oF. Much

attention will be given to the thermal effects on material properties and stress states in

the detail design and analysis phases.

Fuel Storage

61

3.8 Class II Landing Gear

This section discusses the design of landing gear in terms of stroke length, tire

diameter, and strut diameter. The landing gear must be retractable so as not to

interfere with the radar. More importantly, the landing gear must be retractable with

skis or conventional wheels as the Meridian will be operated from snow and paved

runways. The Red, White, and Blue designs all incorporated tricycle type landing

gear that retracted on a tilted pivot into the fuselage. While this is feasible for

conventional tires, this retraction scheme does not work with skis. The tricycle gear

had several other design problems:

• The nose gear had to be mounted far enough from the propeller to

leave room for the nose ski. This required a very wide gear to meet

lateral tipover.

• The Meridian should be able to be shipped in a 20 foot container,

which is approximately 90 inches wide. Lateral tipover requirements

called for a wider gear than this, so the gear would have to be removed

for shipping.

• There were no commercially available landing gear similar to the

previous design.

All of these problems lead to the development of a new landing gear integration

scheme. The gear disposition was changed to a tail dragger to solve the nose ski

integration and lateral tip-over problems. The landing gear were then moved to pods

mounted to the wing. This had two benefits:

62

• The landing gear could be purchased commercially

• The landing gear retract straight aft, which allows for ski retraction

The decision to put the landing gear on the wing calls for the use of either an oleo,

pneumatic, or rubber damped type strut. The oleo gear was chosen as this type of

gear is commercially available. The landing gear selection will be discussed further

in Section 3.8.3.

The following assumptions were made for the landing gear design:

• The main gear shall be able to sustain 100% of the static load. (This is

due to the tail dragger configuration.)

• The gear will be sized for a maximum touchdown rate of 10 ft/s.

• The stroke length will be sized such that a 10 ft/s decent rate imparts

1g on the airframe.

• The strut will have an energy absorption efficiency of 80%.

• The strut will be sized for skis, thereby setting the tire deflection to

zero.

3.8.1 Tire Selection

The main gear tires will be 3.00 x 4 tires. These tires have an outer diameter of 10.0

inches, a width of 3.2 inches, a maximum pressure of 50 psi, and weigh 3.5 lbs each.

The tail wheel tire will be a 6.0 inch diameter solid rubber tire, which weighs 4.75

lbs. As will be discussed in Section 3.8.3, the main and tail gears are commercially

available parts currently used on homebuilt aircraft.

63

3.8.2 Strut Sizing

The strut stroke length and diameter sizing were performed using the methods

described in [6]. The stroke length is calculated by determining the touchdown

energy using Equation 3.4. The gear absorption energy equation is then used to

determine the appropriate stroke length using Equation 3.5. The results are shown in

Table 3.22.

gwWE t

LT

2

21

= Equation 3.4

s

ttgms

T

s

sNPn

E

Sη

η−

=

Equation 3.5

Where:

• Ss = Strut stroke length

• Et = Touchdown energy

• ns = Number of struts

• Pm = Max static load per gear

• Ng = Ratio of max load to static load

• ηt = Tire energy absorption efficiency

• st = Tire deflection

• ηs = Strut energy absorption efficiency

64

Table 3.22 - Landing Gear Strut Sizing

Parameter Units ValueWL lbs 1082wt fps 10ns ~ 2Pm lbs 541Ng g 1ηt ~ 0st in 0ηs ~ 0.6Diameter in 0.1Stroke in 2.6

3.8.3 Landing Gear Integration

The landing gear placement, integration, and sizing were iterated such that a

commercially available landing gear could be integrated with the Meridian. This

greatly increases the feasibility of manufacturing the Meridian in the time allotted as

landing gear development is a fairly complicated process. The landing gear strut

produced for the Lancair Legacy homebuilt aircraft [37] will be used for the main

gear (Figure 3.24 and Figure 3.25).

Figure 3.24 - Lancair Legacy Landing Gear Strut [37]

65

Figure 3.25 - Lancair Legacy Landing Gear Installation [38]

The tail wheel will also be purchased commercially. The tail wheel assembly is

manufactured by Matco [39] and is commonly used on homebuilt aircraft.

Figure 3.26 - Matco Tailwheel Assembly [39]

66

3.9 Structural Arrangement

The purpose of this section is to discuss the proposed structural arrangement for the

Meridian UAV. This includes material selection, structural layouts for the wing, v-

tail, and fuselage, as well as preliminary structural sizing.

The mission of the Meridian is considered to be extreme in that the locations of

operation, Greenland and Antarctica, are known for extremely cold weather. While

this must be considered during the structural arrangement and material selection, it

must be noted that the temperatures the Meridian will experience are not much

different from High Altitude Long Endurance (HALE) UAVs. In fact, the Meridian

mission may be less extreme than a HALE UAV because it will not experience large

changes in temperature throughout a flight. This is mentioned only to emphasize the

fact that the material selections should not be arbitrarily limited due to cold weather

operations. Rather, the changes in material properties due to temperature should be

acknowledged and accounted for in the design process such that the final product is

an optimized solution in terms of weight, manufacturability, and service life.

One of the primary drivers of the material selection and structural layout is the

advanced development time requirement. For the Meridian to be a successful project,

manufacturability has to play a big role in the structural design process. In addition,

many of the structures will be manufactured and assembled by graduate students with

limited manufacturing experience. Therefore, the aircraft should be designed in such

a way that limits the manufacturing skills and facilities required as is often done with

homebuilt aircraft. These two concerns warrant the need for a limited part count as

67

well as a high level of automated processes such as computer numerically controlled

machining.

3.9.1 Wing Structure

The structural arrangement of the wing for the Meridian was driven by the

following concerns and requirements:

• Shipping requirement

• Hard point mounting requirement for antennas

• Fuel system integration

• Cost

• Manufacturability

• Weight

• On site storage facility limitations

The shipping, storage, and hard point requirements were determined to be the most

critical and therefore had the biggest effect on the wing structural arrangement. The

shipping and storage limitations [5] were such that the wing had to be designed in at

least three pieces. In terms of structural optimization, the best solution would be to

make the wing in two pieces. This however, does not consider the other system

requirements and limitations such as landing gear and fuel tank integration nor does it

consider manufacturability.

The structural layout of the wing was determined by integrating the landing gear

placement, fuel tank sizing, control surface sizing, shipping requirements, and

manufacturing limitations. The final solution is a three-piece wing: the inboard

68

section contains the fuel tanks and landing gear and the outboard sections contain all

of the control surfaces. The outboard sections are removable for shipping and storage

in small field hangars. In addition the length of the longest part in the wing is less

than that of the current composite curing facilities at the University of Kansas (~10

ft).

Several possible arrangements were investigated for the wing structure including:

• Single spar

• Two Spar

• Three Spar

• Tube Spar

The single spar concept was eliminated as the control surfaces will require some

sort of closeout mechanism. The three spar concept was eliminated based on

preliminary structural sizing analyses. The two spar concept was selected as the

primary configuration with the tube spar as a secondary option. The wing is designed

with a rectangular forward spar and a c-channel rear spar. The spar of the outboard

wing slides into the inboard spar and is held by fasteners on top and bottom. This

allows the outer portion of the wing to be removable without adding a great deal of

complexity of weight.

69

Figure 3.27 - Wing Structural Layout

The material selection for the wing was influenced primary by manufacturability,

load types, and thermal considerations. In terms of manufacturability, composite

skins allow for a high level of automation in the tooling manufacturing and provide

excellent surface finish. In terms of the substructure, there are several locations

where loaded fasteners are required such as the landing gear attachment, wing joint,

and antenna hard points. This warrants the use of aluminum in several of the

structural components such as the forward and aft spar as well as several of the ribs.

The combination of different materials in the wing has implications in terms of

thermal expansion. These will be investigated further in the detailed design and

analysis of the structure.

70

3.9.2 Fuselage Structural Layout

The structural layout of the fuselage was driven by the following:

• Wing-Fuselage Integration

• Manufacturability

• Weight

• Engine Installation

• Payload Integration

• Accessibility Requirements

The structural layout of the fuselage was integrated with the configuration design in

terms of wing and payload placement. The wing placement, which is driven by the

aircraft center of gravity was iterated until the main spar of the wing was collocated

with the firewall. To produce structurally efficient aircraft designs, this type of

synergy must be implemented between the design aspects such that the amount of

structural members required is decreased. By locating the landing gear, wing main

spar, and fuselage firewall at the same fuselage station, the amount of heavy

structural members has been decreased, which provides weight savings and improves

the manufacturability of the vehicle.

The primary frames in the fuselage were placed at the locations of the wing spars,

payload hatch closure, payload rack mount, fuselage split, and v-tail spars. The

remainder of the fuselage frames were spaced according to preliminary buckling

calculations. The upper longerons were placed in line with the top engine mounting

bracket as well as the payload hatch opening. The lower longerons were located at

71

the upper surface of the wing and coincide with the lower engine mounting brackets.

Two frames were located at the forward and aft v-tail spar locations. These frames

were also used to mount the tailwheel assembly.

Figure 3.28 - Fuselage Structural Layout

72

Figure 3.29 - Wing-Fuselage Attachment

The aircraft structure was iterated several times such that the fuselage and center

wing section would fit in a standard 20 foot container [9]. The goal was to minimize

the amount assembly that would have to be performed on-site. This is important for

shipping, but also for on-site storage. The projected hangar size is approximately 15

feet wide, which means the wings must be removed after every flight. The aircraft is

shown in a 20 foot container in Figure 3.30.

73

Figure 3.30 - Standard 20 Foot Shipping Container Door [9]

The engine mount will be procured from the engine manufacturer and will be very

similar to the mount shown in Figure 3.31.

74

Figure 3.31 - Typical Engine Mount for the Innodyn 165TE

3.10 Manufacturing Breakdown

The aircraft skin is divided into 6 pieces: 2 for the cowl, left and right sections for

the forward fuselage, and top and bottom sections for the aft fuselage sections.

Again, manufacturability was a primary driver in the fuselage design as the leading

edge of the v-tail was placed such that it coincides with the mold line of the fuselage.

This allows for the aft fuselage and v-tail skin to be continuous, which improves

structural rigidity and reduces the parts count.

75

Figure 3.32 - Manufacturing Breakdown

3.11 Cost Analysis

The cost estimate for the Meridian UAV was created using the AAA software [7]

and the methods described in [6]. The vehicle cost is broken down into research,

development, test, and evaluation or RDT&E costs; and acquisition cost. The

methods of [6] are typically used for production type aircraft that will be sold for

some profit. This Meridian is strictly a research aircraft developed for a specific

mission. The marketability of the Meridian, while may be exploited at a later date, is

not part of this cost estimate. For this reason, the cost estimates of [6] were

augmented with quotes from vendors for items such as avionics, tooling, and engines.

76

The first step in the cost estimation is to determine the Aeronautical Manufacturer’s

Planning Report (AMPR) weight. This is defined as the vehicle empty weight less all

of the items that will be purchased from vendors such as the engine, actuators,

avionics, wheels, etc. The AMPR weight of the Meridian was estimated to be:

• WAMPR = 260 lbs

The next step in the cost estimation was to develop the hourly rates to apply to each

cost estimate. This project is different from a typical aircraft manufacturing process

as much of the work will be performed by students, whom work at much lower rates

than typical industry standards. Average rates for manufacturing and engineering

time were developed based on the rates for undergraduate students, graduate students,

professors, and industry labor as shown in Table 3.23. The expected breakdown of

time is also shown in Table 3.23, which was used to create a time-weighted average.

The industry rate was included in the wage calculations as some of the part

manufacturing will be outsourced. The rates shown in Table 3.23 include typical

overhead rates.

Table 3.23 - Engineering and Manufacturing Rate Estimation

% of Total Time Hourly Rate % of Total Time Hourly Rate% $/hr (2006) % $/hr (2006)

Undergraduate 15 $16.00 30 $16.00Graduate 60 $24.00 60 $24.00Professor 15 $96.00 0 $96.00Industry 10 $60.00 10 $60.00Total (Averged) $37.20 $25.20

Engineering Labor Manufacturing Labor

77

3.11.1 Research, Development, Test, and Evaluation Costs

The total RDT&E cost for an aircraft is defined as the sum of the airframe

engineering and design cost; development, support and testing cost; flight test

airplane cost; and flight test operations cost. This cost is then adjusted by factors

accounting for test facilities, profit, and financing. This was not done for the

Meridian, however, as there will be no profit or financing, and the manufacturing

facilities will be paid for by other university funding.

The following assumptions were made in the RDT&E cost estimation:

• The number of aircraft built during the RDT&E phase is 1.

• The workforce is assumed to be relatively skilled in Computer Aided

Design

• The engine cost was estimated at $30,000 per the manufacturer.

• The propeller cost was assumed to be $5,000 per the manufacturer.

• The avionics cost was estimated at $100,000 per the manufacturers.

• No profit or financing were included in the RDT&E phase.

The total cost for the RDT&E phase was determined to be $2.018 million. The cost

breakdown is shown in Table 3.24 on page 79.

3.11.2 Acquisition Cost

The acquisition cost of an aircraft is defined as the sum of the manufacturing cost

and the profit. As there will be no profit for the Hawkeye, this reduces to simply the

manufacturing cost, which is comprised of the airframe engineering and design cost

for the production phase; the airplane program production cost; and the production

78

flight test operations cost. The following assumptions were made for the acquisition

cost estimation:

• The total number of aircraft produced for the production phase is 1.

• The manufacturing rate is assumed to be 0.1 aircraft.

• The interior costs were set as $0.

• 40 hours of flight testing at $500/hr with an overhead factor of 4.0

were assumed for the production vehicle.

• No profit or financing were included in the production cost estimate.

The total acquisition cost was estimated as $1.009 million. The cost breakdown for

the acquisition phase is shown in Table 3.24.

3.11.3 Cost Estimate Summary

The RDT&E and acquisition cost estimates are summarized in Table 3.24. The

total costs are broken down into overall categories in Table 3.25 and by RDT&E and

production categories in Table 3.26.

79

Table 3.24 - RDT&E and Acquisition Cost Summary

Item Cost10$

RDTEEngineering and Design 0.232Development, Support, and Testing 0.087RDTE Labor Costs 0.872Material Costs 0.379Avionics Equipment 0.1Tooling 0.15Quality Control 0.113Engine 0.035Flight Test Operations 0.05

2.018Production Cost

Airframe Engineering and Design 0.031Labor 0.386Production Materials 0.277Production Avionics 0.15Manufacturing Tooling 0Manufacturing Quality Control 0.05Engines 0.035Flight Test Operations 0.08

1.009

Acquisition Cost (2 Vehicles) 3.027

Table 3.25 - Cost Breakdown by Overall Category

Item Cost10$

Labor 1.608Materials 0.656Avionics 0.25Tooling 0.15Quality Control 0.163Engines 0.07Flight Test Operations 0.13Total 3.027

80

Table 3.26 - Cost Breakdown by RDT&E and Production Categories Item Cost

10$

RDTEEngineering and Design 0.232Development, Support, and Testing 0.087Test Aircraft 1.649Flight Test Operations 0.05

2.018Production Cost

Airframe Engineering and Design 0.031Production Manufacturing 0.898Flight Test Operations 0.08

1.009Acquisition Cost (2 Vehicles) 3.027

Labor54%

Materials22%

Avionics8%

Tooling5%

Quality Control5%

Engines2%

Flight Test Operations4%

Figure 3.33 - Cost Breakdown by Overall Category

81

3.11.4 Cost Estimate Justification

The cost estimate performed is based on conventional aircraft production costs.

The viability of using these methods for a UAV is questionable. Therefore, the

estimated cost was tabulated against several current UAV systems for comparison as

shown in Table 3.27. It is important to note that the cost of these systems is listed in

terms of the vehicle costs and the system costs, which include ground support

equipment and a certain number of vehicles. It is difficult to estimate the cost of one

vehicle with ground support equipment, therefore the system cost was divided by the

number of aircraft per system. This gives a more reasonable estimate as to the actual

cost of functional UAV. The aircraft cost and the cost per aircraft based on system

cost were plotted versus payload weight in Figure 3.34 and Figure 3.35 respectively.

Figure 3.34 shows that the estimate vehicle cost of the Meridian is almost exactly on

the linear regression. This indicates that the vehicle cost estimate is reasonable. The

cost per aircraft based on system cost plotted in Figure 3.35 shows how the Meridian

is a more cost-effective system because the ground support equipment (ground

station, charging system, etc) is already included in the vehicle cost estimate. (The

ground station costs are included in the avionics cost estimates.)

Table 3.27 - Current UAV Procurement Cost [40] System Aircraft Weight Payload Aircraft Cost System Cost Number Cost Per Aircraft

lbs lbs FY06$ mil FY06$ mil Acft/System In System FY06$ milDragon Eye 4 1 0.03 0.14 3 0.05RQ-7A Shadow 216 60 0.41 13.24 4 3.31RQ-2B Pioneer 307 75 0.68 17.93 5 3.59RQ-8B Fire Scout 1,765 600 4.27 22.83 4 5.71RQ-5A Hunter 1,170 200 1.25 27.62 8 3.45MQ-1B Predator 1,680 450 2.81 25.75 4 6.44MQ-9A Predator 3,050 750 5.42 47.01 4 11.75RQ-4 (Block 10) Global Hawk 9,200 1,950 19.81 60.15 1 60.15RQ-4 (Block 20) Global Hawk 15,400 3,000 27.62 64.84 1 64.84Meridian 1,082 165 1.51 3.03 2 1.51

82

Dragon Eye

Shadow

Pioneer

Hunter

PredatorFire Scout

Predator

Global HawkGlobal Hawk

Cost(FY06$ 10^6) = (0.0089)WPL

0.00

0.01

0.10

1.00

10.00

100.00

1 10 100 1,000 10,000Payload Weight, lbs

Cos

t (FY

06$

10^6

)

Meridian

Figure 3.34 - UAV Cost in Terms of Payload Weight

83

Global HawkGlobal Hawk

Predator

Fire Scout

Predator

HunterPioneerShadow

Dragon Eye

Cost (FY06$ 10^6) = (0.0225)WPL

0.01

0.10

1.00

10.00

100.00

1 10 100 1,000 10,000Payload Weight, lbs

Cos

t (FY

06$

10^6

)

Meridian

Figure 3.35 - UAV Cost Based on System Cost Versus Payload Weight

84

4 Conclusions This document summarizes the redesign of the Meridian UAV based on the

response from the Preliminary Design Review. The antenna type has been changed

from a bow-tie to a Vivaldi or exponential antenna. In addition, the aircraft has been

modified to incorporate several commercially available off-the-shelf parts. The goal

with the redesign of the Meridian is to produce a design that is not only novel, but is

feasible considering the extremely short development time. This meant integrating

manufacturability, performance, and operational constraints into the design process.

The product is a vehicle that can be shipped in a standard 20 foot container and

quickly assembled and disassembled with minimal tools. The Meridian is the

smallest turboprop powered UAV in the world. It is also one of the only UAVs with

retractable ski landing gear. The purpose of this continued development of the

Meridian is to completely flush out the ‘best’ new UAV design based on the mission

specification.

85

5 References 1 www.worldaerodata.com/countries/Antarctica.php 2 www.is.northropgrumman.com 3 www.uavforum.com 4 www.uav.com 5 “Mission Concepts for Uninhabited Aerial Vehicles in Cryospheric Science

Applications”. University of Kansas Remote Sensing Laboratory. KS, 2004. 6 Roskam, Jan. Airplane Design: Parts I-VIII. DARCorporation. Lawrence, KS.

1997. 7 Advanced Aircraft Analysis Software. DARCorporation. Lawrence, KS. 2005. 8 Simons, Martin. Model Aircraft Aerodynamics, 4th Edition. Nexus Special

Interests, 1999. 9 www.oceanairlogistics.com 10 www.piaggioamerica.com/ 11 Raymer, Daniel P. Enhancing Aircraft Conceptual Design using

Multidisciplinary Optimization. Ph. D. Thesis, Swedish Royal Institute of Technology, Stockholm, Sweden, 2002.

12 Worldwide UAV Roundup. www.aiaa.org/images/PDF/WilsonChart.pdf. 2003. 13 Munson, Kenneth. Jane’s Unmanned Aerial Vehicles and Targets Issue 11. 1999. 14 Lambert, Mark ed. Jane’s All the World’s Aircraft 1993-94. Jane’s Information

Group. Alexandria, VA. 1993. 15 Sobieszczanski-Sobieski, J., “Multidisciplinary Design Optimization: An

Emerging New Engineering Discipline,” Advances in Structural Optimization (483-496), Kluwer Academic Publishers, the Netherlands, 1995.

16 www.airliners.net. April 25, 2006. 17 www.aviataircraft.com. April 25, 2006. 18 www.staggerwing.com. April 25, 2006. 19 www.comnap.com. May 19, 2005 20 http://www.scoop.co.nz/stories/PO0601/S00042.htm. May 19, 2006. 21 www.aaicorp.com. May 19, 2006. 22 www.genaero.com. May 19, 2006. 23 www.diamondair.com. May 19, 2006. 24 http://www.diamond-air.at/en/press/pressarchive/40820.htm. May 19, 2006. 25 Donovan, William. “CReSiS Airborne Platform Summary”. The University of

Kansas. 2006. 26 www.rotax-aircraft-engines.com. May 19, 2006. 27 Hoerner, Sighard. Fluid-Dynamic Drag. Published by Author. Great Britain,

1958. 28 Von Mises, Richard. Theory of Flight. Dover Publications. New York, 1959. 29 Barrett, Ron. “Discussion Regarding Biplane Design”. The University of

Kansas. May 10, 2006.

86

30 Allen, Christopher. “Discussion Regarding Antenna Design.” The University of Kansas. April 2006.

31 Palais, July, et al. “Meeting with NSF Representatives.” February, 2006. 32 Munk, Max M. “General Biplane Theory.” NACA Report No. 151. 1923. 33 www.nsf.gov. May 19, 2006. 34 “Science Requirements for Field Work in CReSIS. The University of Kansas.

September 15, 2005. 35 www.faa.gov. August 11, 2006. 36 www.cloudcaptech.com. August 11, 2006. 37 www.lancair.com. August 11, 2006. 38 http://www.lancairlegacy.com/links.html. August 11, 2006. 39 www.aircraftspruce.com. August 11, 2006. 40 Cambone, Stephen, et al. “Unmanned Aircraft Systems Roadmap 2005-2030.”

Office of the Secretary of Defense. August 2005.

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