Usc scad ancile

Southern California Aircraft Design

Ancile June 14, 2013

Executive Summary The ever evolving threats, which the United States experiences from areas of the world such as Iran and North Korea, have increased calls for varying forms of ballistic missile defense. Yet this comes on the heels of the cancellation of the Airborne Laser Program (ABL) in 2010 after the ABL flirted with success.

Since the ABL costs more than $100,000 per hour [1], a future replacement must be able to operate successful while being significantly cheaper to operate. This feature is crucial for an ABL replacement as it would need to be up in the air for an extended time period of up to 24 hours. To reduce costs during extended periods, it is paramount that the replacement be unmanned. This would allow the UAS to be lighter and more efficient than a manned counterpart. Additionally, an unmanned system is not as susceptible to pilot fatigue, thus reducing the probability of a mistake being made. The need also remains for the aircraft, for such a replacement to fly above adverse weather to avoid adverse performance of the laser.

SCAD believes that Ancile answers the need for theater ballistic missile defense. The focus of Ancile is to provide a low-cost, long endurance UAS system that can destroy ballistic missile while they are in their boost phase. Since Ancile uses a solid state laser, the use of sophisticated cooling systems efficiently manages the heat dissipated from firing the laser. Because the cooling system is so effective, the laser could be fired for a longer sustained burst than specified in the request for proposal. The limiting factor on the rate of fire of the laser is the time it takes to recharge the system that provides power to the laser. Ancile also utilizes high aspect ratio wings to improve loiter performance to succeed in its primary mission of providing defense from ballistic missiles. The advanced communications suite on the Svallinn allows it to be controlled from a control station anywhere on the globe regardless of the aircraft’s location relative to that control station.

With an endurance of 24 hours and a maximum unit cost of approximately $82 million, SCAD stands behind Ancile as a low cost solution to U.S, and possibly NATO, missile defense needs.

Acknowledgements We would like to acknowledge the contributions by a former student who lent his support in completing this proposal, Geoff Legg. Additionally we would like to thank Dr. Geoffrey Spedding for providing guidance and moral support in establishing and sustaining this year’s team and for lending the resources of the Aerospace and Mechanical Engineering Department’s resources so that we may procure software and texts that helped in designing the Ancile and writing this proposal. We would also like to thank Vivek Mukhopadhyay, from NASA Langley, for his provision of the preliminary flutter software program.

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Andrew Levinson Member # 457142 USC Lead Engineer Configuration

Keith Holmlund Member # 413045 USC Aerodynamics Campus Project Manager

Controls Dan Eusebio Member #473412 UCLA Laser Systems Power Storage and Distribution

Patrick Gendotti Member # 306354 USC Avionics Systems Architecture

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James Wong Member # 481174 USC Cooling Systems Aircraft Materials

Azadeh Keyvani Member # 280596 USC Weight and Balance Engineering

Darin Gaytan Member # 279164 USC CAD Integrator

Herbert Turner Member # 480164 USC Aircraft Performance High-Lift Devices

Richard Boles Member # 437301 USC Propulsion

Jen Sheriff Member # 476611 UCLA Avionics Power Systems

Dr. Geoff Spedding Member # 233796 USC Chairman & Project Advisor

Dr. Charles A. Radovich Member # 189769 USC Faculty Advisor

Contents Executive Summary ................................................................................................................................ 1 Acknowledgements ............................................................................................................................... 1 Table of Figures ........................................................................................................................................... 6 Nomenclature ......................................................................................................................................... 8 Acronyms ............................................................................................................................................... 11 1) Introduction ....................................................................................................................................... 13 1.1) Request for Proposal Interpretation ....................................................................................... 13 1.2) Requirements Matrix ................................................................................................................. 13 1.3) Design Process and Methodology ......................................................................................... 14 2) Mission Profile .................................................................................................................................... 14 2.1) Primary Mission Profile ............................................................................................................... 14 3) Preliminary Sizing .............................................................................................................................. 15 3.1) Mission Analysis .......................................................................................................................... 15 3.2) Preliminary Drag Polars ............................................................................................................. 16 3.3) Constraint Diagram................................................................................................................... 17 4) Configuration Description ....................................................................................................................... 17 4.1) Overall ............................................................................................................................................ 17 4.2) Wing ............................................................................................................................................... 18 4.2.1)

Landing Gear................................................................................................................. 20

4.2.2)

Aspect Ratio Justification ............................................................................................ 20

4.3) Fuselage...................................................................................................................................... 23 4.4) Canard and Vertical Tail .......................................................................................................... 24 4.5) Propulsion ...................................................................................................................................... 25 5) Aerodynamics ........................................................................................................................................ 26 5.1) Airfoil Selection .......................................................................................................................... 26 5.2) High Lift Devices ........................................................................................................................ 26 5.3) Detailed Drag Polars ................................................................................................................. 27 5.4) Computational Fluid Dynamics ......................................................................................................... 28 5.4.1)

Airfoil Validation ............................................................................................................ 28

5.4.2)

Canard Study................................................................................................................. 29

5.4.3)

Analysis of Final Aircraft Configuration ...................................................................... 30 2|Page

6) Propulsion .......................................................................................................................................... 31 6.1) Turbofan versus Turboprop ...................................................................................................... 31 6.2) Engine Selection and Characteristics .................................................................................... 32 6.3) Engine Inlet Area ....................................................................................................................... 34 7) Avionics .............................................................................................................................................. 35 7.1) Introduction ................................................................................................................................ 35 7.2) Requirements ............................................................................................................................. 36 7.3) Communication ........................................................................................................................ 39 7.4) Sensors ......................................................................................................................................... 40 7.5) Flight Control Computers ......................................................................................................... 41 7.6) Power .......................................................................................................................................... 42 7.7) System Architecture .................................................................................................................. 43 8) Laser System ...................................................................................................................................... 44 8.1) Power .......................................................................................................................................... 44 8.2) Cooling........................................................................................................................................ 47 8.2.1)

Background ................................................................................................................... 47

8.2.2)

Laser Housing Sizing ...................................................................................................... 49

8.2.3)

Channel Sizing ............................................................................................................... 49

8.2.4)

Fluid Dynamics Simulations .......................................................................................... 50

8.2.5)

Passive mixing using ramps .......................................................................................... 50

8.2.6)

Routing through multiple channels ............................................................................ 53

8.2.7)

Comparison of channel geometries .......................................................................... 53

8.2.8)

Placement and Configuration.................................................................................... 54

8.2.9)

Power Requirements..................................................................................................... 54

8.3) Layout ......................................................................................................................................... 56 9) Materials Selection ........................................................................................................................... 56 9.1) Structure: 2024 – T3 (Flat sheet) Aluminum ............................................................................ 56 9.2) Cooling channels: 6063-T5 Aluminum .................................................................................... 56 9.3) Cooling Liquid: JP-8................................................................................................................... 57 9.4) Nozzle: Grade 6 Ti Alloy ............................................................................................................ 59 9.5) Wings and skin: Woven AS4 Carbon Fiber ............................................................................ 60 9.6) Leading Edge of Wings: 2124 – T851 ...................................................................................... 60 9.7) Radar Dome: Fiber Glass ......................................................................................................... 60 9.8) Laser Housing: Duroplast Thermoset....................................................................................... 60 3|Page

10) Dynamics and Controls ......................................................................................................................... 60 10.1) Un-augmented Dynamics ............................................................................................................... 60 10.2) Mil-Std Requirements .................................................................................................................... 64 10.2.1 Longitudinal Dynamic Stability ........................................................................................ 64 10.2.2 Lateral Dynamic Stability .................................................................................................. 66 10.3) Augmented Dynamics .................................................................................................................... 69 11) Weight Justification .............................................................................................................................. 74 11.1) Detailed Weight Analysis ................................................................................................................ 74 11.2) Class III Weight Estimation ............................................................................................................. 75 11.3) Center of Gravity Analysis .............................................................................................................. 75 12) Performance ................................................................................................................................... 77 12.1) Takeoff Performance .............................................................................................................. 77 12.2) Climb Performance ................................................................................................................ 78 12.3) Loiter Performance ................................................................................................................. 79 12.4) Landing Performance ............................................................................................................ 81 13) Recommendations ........................................................................................................................ 82 13.1) Design Recommendations .................................................................................................... 82 13.1.1)

Aerodynamic and Configuration Recommendations ........................................... 82

13.1.2 Avionics Recommendations ............................................................................................ 82 13.2) Capability Upgrades .............................................................................................................. 83 13.2.1)

Avionics Upgrades ........................................................................................................ 83

13.2.2)

Laser System Upgrades ................................................................................................ 83

14) Cost .................................................................................................................................................. 84 15.0 Software Used Description .......................................................................................................... 86 15.1 Advanced Aircraft Analysis (AAA) ........................................................................................ 86 15.2 Star CCM+ v8 ............................................................................................................................ 86 15.3 GasTurb 11 ................................................................................................................................. 86 15.4 Conceptual Flutter Analysis of Cantilever Wing (CFACW) ................................................ 86 15.5 DesignFOIL ................................................................................................................................. 87 15.6 Ansys CFX ................................................................................................................................... 87 15.7 Solidworks ................................................................................................................................... 87 15.8 Rhinoceros w/ VRAY ................................................................................................................. 87 15.9 Microsoft Excel .......................................................................................................................... 87

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16)Works Cited ...................................................................................................................................... 87

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Table of Figures Figure 2.1.1: The mission profile of the UAS: 1 – Takeoff 2 – Climb 3 – Cruise 4 – Loiter 5 – Return to Cruise 6 – Descent 7 – Landing ......................................................................................... 14 Figure 3.2.1: Results from preliminary aerodynamic projections. (a) Preliminary drag polars for different aircraft mission segments. (b) Parametric analysis of lift and drag data. C_L corresponding to the maximum C_Lâ „C_D at constant Mach number maximizing the range. đ?‘Şđ?‘ł corresponding to the maximum đ?‘Şđ?‘łđ?&#x;Ž. đ?&#x;“đ?‘Şđ?‘Ť maximizes the range at a constant altitude. The parameter đ?‘Şđ?‘łđ?‘Şđ?‘Ťđ?&#x;‘đ?&#x;? maximizes the SAR of the aircraft and was selected based on the recommendation given in ESDU 73019 as a measure of merit, defining a design region for the cruise đ?‘Şđ?‘ł of the aircraft................................................................................................................. 16 Figure 3.3.1: Constraint Diagram ........................................................................................................ 17 Figure 4.2.1: Wing Planform and Dimensions in feet ....................................................................... 19 Figure 4.2.2: Wing Structure and Locations of Fuel Tank (Green) and Flaps (Blue) ................... 20 Figure 4.2.3: Flutter and Regier number boundaries as a function of Mach number for low sweep angles ........................................................................................................................................ 22 Figure 4.2.4: Summary table with flutter and dynamic pressure boundaries as a function of Mach number. The two x show the dynamic pressure of the aircraft in it’s loiter and cruise conditions respectively at an altitude of 40,000 ft. ......................................................................... 22 Figure 4.3.1: Fuselage Dimensions...................................................................................................... 23 Figure 4.3.2: Fuselage Structure .......................................................................................................... 24 Figure 4.4.1: Vertical Tail and Canard Structure .............................................................................. 24 Figure 4.4.1: Vertical Tail and Canard Planforms ............................................................................ 25 Figure 5.2.1: Maximum Lift Coefficient vs. Flap Deflection at for several flap chord to wing chord ratios............................................................................................................................................ 26 Figure 5.3.1: Drag Polars at key mission segments. Cruise and Loiter points depict optimal flight performance. Takeoff and Landing points indicate takeoff and landing at a CLmax of 2.2 and 2.7 respectively ....................................................................................................................... 27 Figure 5.3.2: Drag Breakdown at Cruise. The 12% of drag generated by the fuselage shows the successfully implantation of the laser with respect to the performance of the aircraft ... 27 Figure 5.4.1: Example CFD Cell Configuration ................................................................................. 28 Figure 5.4.2: Canard Below Wing CFD Result ................................................................................... 29 Figure 5.4.3: Wing CFD Analysis .......................................................................................................... 30 Figure 5.4.4: Sample CFD Results ........................................................................................................ 31 Figure 6.2.1: Carpet Plot of SFC as a Function of Thrust ................................................................. 33 Figure 7.1.1: The spaced distribution of the computer resources aboard an aircraft. This distribution prevents catastrophic failures if one or more computer units are destroyed or malfunction. .......................................................................................................................................... 35 Figure 7.2.1: The front view of the fuselage contains the avionics placed towards the front of the aircraft, especially within the fuselage “bubble.â€? ................................................................... 38 Figure 7.2.2: The rear view of the fuselage contains the avionics placed towards the aft end of the aircraft. These avionics mostly regulate the engine or are “distributedâ€? to improve redundancy, reliability, and durability. ............................................................................................. 38

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Figure 7.2.3: The overall avionics placement throughout the fuselage is depicted here. This image contains the avionics described in the Figures 2 and 3 above, with an additional GPS unit and IFF unit towards the middle of the fuselage. .................................................................... 39 Figure 7.3.1: This diagram depicts the different communications links between a sample UAV (Global Hawk) and support infrastructure and command centers. ........................................... 39 Figure 7.4.1: Capabilities of the AN/APG-77 Synthetic Aperture Radar System ........................ 40 Figure 7.5.1: Overview of Systems Architecture for the Control Electronics (ICP, ECC, PMC, PIU) .......................................................................................................................................................... 41 Figure 7.6.1: Hierarchy of power distribution. ................................................................................... 42 Figure 7.7.1: Specific Avionics Systems Architecture for the Major Components of the UAV (neglecting the electronic controller architecture depicted in Figure 7 earlier) ...................... 43 Figure 8.1.1: Shows the Hierarchy of the Power System where B1-B8 denotes the battery packs and 7C and 6C denote the number of super capacitors in that channel. ................... 45 Figure 8.1.2: Shows the charging circuit for each block of 7 super capacitors. The ESR for each S.Capacitor is also included. ................................................................................................... 45 Figure 8.1.3: Shows the Voltage output of each capacitor with respect to time in the 7 capacitor block circuit. ....................................................................................................................... 46 Figure 8.1.4: Shows the current output across each capacitor and the total current across the resistance with respect to time in the 7 capacitor block circuit. .......................................... 46 Figure 8.1.5: Shows the voltage output of each capacitor with respect to time for the 6 S. capacitor block. ................................................................................................................................... 47 Figure 8.2.1: Number of Channels vs. RE ........................................................................................... 49 Figure 8.2.2: Temperature vs time ...................................................................................................... 50 Figure 8.2.3: Temperature Gradient .................................................................................................. 51 Figure 8.2.4: Streamlines ...................................................................................................................... 51 Figure 8.2.5: side view of the ramps .................................................................................................. 52 Figure 8.2.6: top view of the ramps ................................................................................................... 52 Figure 8.2.7: Temperature Gradient with ramps .............................................................................. 52 Figure 8.2.8: the model for the cooling channels. The fluid entrances and exits are highlighted in blue ................................................................................................................................ 53 Figure 8.2.9: Isometric view ................................................................................................................. 54 Figure 9.3.1: Reynolds Number as a Function of Temperature ..................................................... 57 Figure 9.3.2: Conductive Heat Transfer Coefficient vs. Temperature .......................................... 58 Figure 9.3.3: Prandlt Number as a Function of Temperature......................................................... 59 Figure 10.1.1: Vertical Velocity Throttle Response. .......................................................................... 62 Figure 10.2.1:Phugoid Requirements for Damping Ratio (from MIL-F-8785C)............................. 64 Figure 10.2.2: Short-period frequency requirements (Category B Flight Phases – cruise and loiter). Red squares denote the un-augmented short-period modes, and orange squares denote the augmented short-period nodes. .................................................................................. 65 Figure 10.2.3: Short-period damping ratio requirements (Category B Flight Phases due to loiter and cruise segments being examined) .................................................................................. 66 Figure 10.2.4: Dutch Roll damping ratio, frequency, and product of damping ratio and frequency requirements (Category B loiter and cruise Flight Phases, at Level 1 flight conditions) ............................................................................................................................................. 67

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Figure 10.2.5: Roll mode time constant requirements (Category B Flight Phases given loiter and cruise) ............................................................................................................................................. 68 Figure 10.2.1: Spiral mode time to double amplitude requirements (Category B Flight Phases of loiter and cruise). The first-order time to double (or half) amplitude is derived from the frequency and damping ratio, which give the spiral mode time constant (τ = 1/(Μω)) .......... 68 Figure 10.3.2: State Feedback Control Block Diagram .................................................................. 69 Figure 10.3.3: Z-axis Velocity Throttle Impulse Response ................................................................ 71 Figure 10.3.4: Flight Speed Throttle Response .................................................................................. 72 Figure 10.3.5: Roll Angle Aileron Impulse Response ........................................................................ 73 Figure 10.3.6: Yaw Rate Aileron Impulse Response ......................................................................... 74 Figure 11.3.1: Location of Various Component CGs ...................................................................... 76 Figure 12.1.1: Takeoff Flight Path ........................................................................................................ 77 Figure 12.2.1: ROC vs. velocity at various altitudes including cruise altitude, operational ceiling, and ceiling. .............................................................................................................................. 78 Figure 12.3.1: Elliptical configuration with target outside flight path. Red line denotes portion of flight path where vision angles are insufficient. .......................................................................... 80 Figure 12.3.2: Figure eight configuration with target located along the x axis inside the flight path. Red line denotes portion of flight path where vision angles are insufficient. .................. 80 Figure 12.4.1: Landing Flight Path ...................................................................................................... 81

Nomenclature đ?‘?- Span bht – Span of the horizontal tail bvt – Span of the horizontal tail β – Sideslip angle âˆ†Î˛ – change in sideslip angle ∆đ?›˝Ě‡ – change in first derivative of the sideslip angle đ??śđ??ż – Airplane Coefficient of Lift CLmax – The maximum lift coefficient achieved using high lift devices. đ??śđ??ˇ – Airplane Coefficient of Drag Cp – Pressure Coefficient đ?‘?đ?‘&#x; – Root Chord cđ?‘&#x;â„Žđ?‘Ą – Root Chord of the horizontal Tail cđ?‘&#x;đ?‘Łđ?‘Ą – Root Chord of the vertical Tail đ?‘?đ?‘Ą – Tip Chord cđ?‘Ąâ„Žđ?‘Ą – Tip chord of the horizontal Tail cđ?‘Ąđ?‘Łđ?‘Ą – Tip chord of the vertical Tail 8|Page

đ??ˇ – Drag force or hydraulic Diameter đ?›Ľ – Quarter chord sweep đ?›żđ?‘Ž − Change in aileron deflection Δđ?›żđ?‘Ž − Change in aileron deflection đ?›żđ?‘’ − Elevator Deflection Δđ?›żđ?‘’ − Change in elevator deflection

đ?›żđ?‘&#x; − rudder deflection Δđ?›żđ?‘&#x; − Change in rudder deflection đ?›żđ?‘‡ − Thrust Angle Δđ?›żđ?‘‡ − Change in thrust angle đ??¸đ??´đ?‘™đ?‘œđ?‘? – Location of elastic axis (35% of chord Sec A) Ζd – Dutch Roll Damping ratio Îśp – Phugoid Damping ratio ÎśSP – Short Period Damping ratio đ??š – Flutter number f - Friction factor đ??şđ??˝ – Torsional stiffness ∆θ – Change in the pitch angle ∆đ?œƒĚ‡ – Change in the first derivative of the pitch angle đ??źđ?‘?đ?‘–đ?‘Ąđ?‘?â„Ž – Pitch axis moment of inertia đ??ź60 – Running pitching moment at 60% semi span Ka – Torsional frequency factor đ??ż – Lift force L/D – Lift to Drag Ratio L/DTO – The maximum lift to drag ratio during takeoff ΛLE – Leading Edge Sweep Angle Λđ??żđ??¸â„Žđ?‘Ą – Leading Edge Sweep Angle of the horizontal Tail M - Mach number Mxx – Moment due to “xxâ€? control parameter

Οo – Mass ratio at sea level Ο – Viscosity 9|Page

p – Roll rate Δđ?‘? − Change in the roll rate Δđ?‘?̇ − Change in the first derivative of the roll rate đ?‘ƒđ?‘&#x; − Prandtl number Î TO – Throttle Setting at Takeoff q – Pitch rate Δđ?‘žĚ‡ − Change in the 1st derivate of the pitch rate Δđ?‘ž − Change in pitch rate đ?‘… – Regier number r – Yaw rate Δđ?‘&#x; − Change in the yaw rate Δđ?‘&#x;̇ − Change in the first derivative of the yaw rate RE – Reynold’s Number Ď - Density đ?‘† – Area Ďƒ- Density ratio T – Temperature or Thrust or Period đ?‘‡/đ?‘Š – Thrust to Weight Ratio u – Forward velocity đ?›Ľđ?‘˘ - Change in forward velocity đ?›Ľđ?‘˘Ě‡ - Change in the forward acceleration đ?‘‰đ?‘… – Regier Surface Velocity Index V – Velocity v- Lateral velocity w- Vertical velocity Δđ?‘¤ − change in the vertical velocity Δđ?‘¤Ě‡ − change in the vertical acceleration đ?‘Šđ?‘’đ?‘Ľ – exposed weight per side of the wing đ?‘Š/đ?‘† – Wing Loading φ – Roll angle âˆ†Ď† – Change in the roll angle ∆đ?œ‘̇ - Change in the first derivative of the roll angle 10 | P a g e

xx – Subscript used to denote that a variable is a function of various (xx) parameters Xmgc – Mean geometric Chord Xxx – Forward Acceleration Imparted to the airplane as a result of a unit change in “xxâ€? parameter Yxx - Lateral Acceleration imparted to the airplane as a result of a unit change in “xxâ€? parameter Ymgc – Y coordinate location of the mean geometric chord Zxx – Vertical acceleration imparted to the airplane as a result of a unit change in “xxâ€? parameter đ?œ”đ?›ź – Torsional frequency đ?œ”đ?‘›đ?‘‘ – Dutch Roll frequency

Acronyms AAA – Advanced Aircraft Analysis ABL - Airborne Laser AIAA – American Institute of Aeronautics and Astronautics CFD – Computational Fluid Dynamics CNI – Communications, Navigations, and Intelligence DE- Directed Energy ECC – Engine Control Computer ESDU-Engineering Sciences Data Unit EW – Electronic Warfare FADEC – Full Authority Digital Electronic Engine Control Unit GPS/INS – Global Positioning System/Inertial Navigation System HALE – High Altitude Long Endurance HPC – High Pressure Compressor HPT – High Pressure Turbine ICBM – Intercontinental Ballistic Missiles ICP – Interface Control Processor IFF – Identify Friend or Foe ILS – Instrument Landing System ISR – Intelligence, Surveillance, and Reconnaissance LPC – Low Pressure Compressor LPT – Low Pressure Turbine MATLAB – Matrix Laboratory 11 | P a g e

MIL-STD – Military Standard MIL-F -8785C – Military Specification, Flying Qualities of Piloted Airplanes MLW – Maximum Landing Weight NACA – National Advisory Committee for Aeronautics NATO – North Atlantic Treaty Organization PMC – Payload Management Computer PIU - Payload Interface Unit RFP – Request For Proposal ROC – Rate of Climb RWR – Radar Warning Receiver SAR – Specific Air Range OR Synthetic Aperture Radar SATCOM – Satellite Communications SCC – Stress Corrosion Cracking SFC – Specific Fuel Consumption SST – Shear Stress Transport TOGW – Takeoff Gross Weight UHF – Ultra-High Frequency UAS – Unmanned Aerial System UAV – Unmanned aerial vehicle

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1) Introduction 1.1) Request for Proposal Interpretation The 2012-2013 American Institute of Aeronautics and Astronautics (AIAA) Graduate Team Aircraft Design Request for Proposal (RFP) bids for an unmanned aerial system that is capable of intercepting a ballistic missile in its boost phase utilizing a directed energy pulse with an entry into service by 2035. The RFP states, “… shall loiter on theater for no less than 24 hours … at least 40,000 ft.” “… shall support no less than five consecutive firings … each lasting no less than 10 seconds.” “… must be compatible with available paved runways that are at least 7,000 feet long.”

The requirement for an endurance of at least 24 hours on theater emphasizes loiter performance of the airframe over both cruise and maneuvering performance that would be of paramount importance for a bomber/ transport aircraft and fighter aircraft respectively. To address this requirement a wing aspect ratio of 15 was selected as the design point for the initial sizing. The selection of a high wing aspect ratio reduces the drag due to lift, a key factor in loiter performance. Additionally, to satisfy the altitude requirement, the loiter phase begins at 40,000 feet. Due to the loiter performance being optimal at a constant lift coefficient, the airframe will gain altitude throughout its loiter phase satisfying the requirement throughout the loiter segment.

To provide the power to the laser and support at least five firings of the laser, a combination of ultracapacitors and batteries was selected. This hybrid system combines the energy density of advanced lithium ion batteries that have been widely adopted throughout the electronics industry with the rapid discharge only capable through the use of capacitors. The energy to be used by the laser is stored within the batteries until the laser is discharged, at which time the ultra-capacitors are charged for the next firing.

In order for the system to be supported on 7,000 foot runways, the design implements fowler flaps into the wing to provide the necessary lift performance during take-off and landing. The runway length requirement also dictated the thrust requirement of the selected power plant.

1.2) Requirements Matrix Parameter HALE UAS Takeoff and Landing Range to Theatre Loiter on Theatre Mission Critical Equipment Entry into Service Lifespan Directed Energy Time between Firings firings Theatre Presence Heat Dissipation

Requirement Unmanned ≤ 7,000 ft 2,000 nm ≥ 24 hr & ≥ 40,000 ft DE weapon and related systems 2035 20 years 5 shots @ 10 seconds Minimized each time on call for 3 days Adequate Heat Dissipation

Svallin Unmanned 7,000 ft 2,000 nm 24 hr @ 40,000 ft DE weapon and related systems 2035 20 years 8 shots 10 Minutes 3 airframes @ 24 hr Active cooling each system

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Power Storage Power Discharge

Ample Power Storage Rapid Power Discharge

Lithium ion batteries Super-capacitor

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1.3) Design Process and Methodology The design process utilized in creating the Svalinn primarily follows the method taught by Dom Palumbo at the University of Maryland, College Park. The method provided the framework for the various processes used in creating the conceptual design and preliminary design. To further improve the efficacy of the design process, input from an industry advisor from Northrop Grumman, Barnaby Wainfan, and the methodology of Jan Roskam were synthesized with the over-arching framework. Many of the calculations referenced in this proposal come directly from classes or are derivations of those equations, and thus cannot be directly credited to a specific person or publication.

The first step in the design process was to create a conceptual design on which the preliminary design would be based. During the conceptual design process, the RFP was analyzed for requirements that would affect design decisions, such as the loiter time requirement. Upon completion of the RFP analysis, the mission presented in the RFP was analyzed for fuel consumption so that a preliminary weight of the aircraft could be estimated and calculations to size the wing and engine could be completed. The specifics of this process are covered in Primary Mission Profile. The remainder of the proposal covers the preliminary design of the entire UAS, from the structural design to the control system design. Using parameters determined in the conceptual design, a preliminary model of the physical appearance of the aircraft was created to accommodate the various systems and support the necessary control surfaces. Development of the avionics architecture and the laser’s power and cooling system was done in parallel to the physical model to shorten development time and utilize team members with different skill sets. The team itself was divided into sub-teams to develop their respective systems with autonomy to make major design decisions without consulting the team lead. Any decisions that affected the overall aircraft were made by the team lead. As more information was acquired, the design was iterated upon to reflect the newly acquired information. The resulting design has evolved significantly from its initial design on a sketch of paper.

2) Mission Profile 2.1) Primary Mission Profile To initially size the UAS, the mission profile of Airborne Laser (ABL) was considered. The ABL was a modified Boeing 747 designed to shoot down ballistic missiles while in boost phase. Though it never flew a combat mission, the ABL was designed to loiter within 400km of the target area at an altitude of 40,000 ft.

Also since the ABL was a modified 747, the ABL lacked survivability features such as stealth and other defense mechanisms, making it necessary for the ABL to be protected

Figure 2.1.1: The mission profile of the UAS: 1 – Takeoff 2 – Climb 3 – Cruise 4 – Loiter 5 – Return to Cruise 6 – Descent 7 – Landing

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via other military assets if it was near or over hostile territory. The RFP takes this requirement into account, thus allowing UAS to be sized without taking survivability into account.

The RFP also requires a maximum aircraft range of 2,000 nm while maintaining a minimum loitering time of 24 hours. With a shorter range and longer loiter time than he ABL, the emphasis is placed on the theatre presence rather and deploying from a forward base. Thus with a greater focus on loiter time, the UAS was sized to a loiter time of 24 hours while having two cruise segments of 2,000 nm for deployment and return. The resulting mission profile as shown in Fig 1 guided the initial sizing of the UAS.

3) Preliminary Sizing 3.1) Mission Analysis Using the mission profile presented in section 2.1, statistics for each mission segment were compiled to present a complete picture of the mission that the UAS will be performing. The altitude was determined from mission requirements under cruise and loiter conditions, as well as general requirements including takeoff and landing in which the aircraft is required to clear a 50 ft. obstacle at the end of the runway. The mission profile considered only a sea level airfield as most airfields that the UAS is expected to operate out of are between sea level and 2000 ft. above sea level, a range in which altitude impacts performance negligibly.

Mach numbers were calculated from the optimal wing loading table to minimize fuel burn through ideal conditions. The distance covered and time elapsed for the cruise and loiter sections were determined by multiplying or dividing the given parameter from the mission requirements by the flight speed during that mission segment. For the climb and decent sections, the time elapsed was calculated from a selected rate of climb/descent. To determine the ground covered during that descent, the climb/descent angle was calculated and then the horizontal component of the velocity was multiplied by the time elapsed. The calculated statistics are presented in Table 3.1.1.

Table 3.1.1: Mission Analysis Statistics

Mission Segment Start up and Takeoff Climb Cruise Loiter Cruise Descent Land and Shutdown

Altitude (ft) 0-50

Mach 0

Distance (nm) 1

Time (min) 5

SFC (lb/lb-hr) 0.32

Weight Fraction 0.975

50-40,000 40,000 40,000 40,000 40,000-50 50-0

0.560 0.722 0.560 0.722 0.2 0

98.1 2000 7710.6 2000 95.4 1

20 289.7 1440 289.7 50 5

0.32 0.5645 0.5336 0.5645 0.32 0.32

0.9959 0.8986 0.6025 0.8986 0.9897 0.992

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3.2) Preliminary Drag Polars Using second order regression analysis presented by Roskam [1], in addition to the results from Svalinn preliminary mission and weight analyses, initial empirical drag polars were generated in order to complete preliminary performance sizing of the aircraft. ESDU Performance Data Item 73019 [3] was consulted in order to choose the most influential and critical parameters affecting fuel burn, range, and endurance. These three parameters were chosen to optimize the lift coefficient region for the aircraft in a cruise flight condition. ESDU 73019 suggest for a maximum Specific Air Range (SAR) by having a maximum value of đ??śđ??ż /đ??śđ??ˇ 3â „2 . SAR represents the sensitivity of the range of the aircraft to takeoff gross weight and, therefore, the amount of fuel burned at a constant altitude. The đ??śđ??ż that maximizes the range at a constant altitude can be found by maximizing đ??śđ??ż 0.5 /đ??śđ??ˇ . As seen in Fig 1a, the SAR is maximized if the aircraft operates at a lift coefficient of .96, which is significantly lower than the lift coefficient of maximum L/D (1.35). Since it can be observed, however, that the đ??śđ??ż â „đ??śđ??ˇ curve in Fig 1b is relatively flat at a lift coefficient value of 1.3, therefore, the optimizing of the aircraft for maximum SAR causes minimal reduction in maximum L/D. Hence to optimize the aircraft at a specific range, the cruise lift coefficient would be .79, the lowest of all cruise C L limits. Therefore, the optimal region can be defined for the optimal cruise CL of the aircraft, allowing the UAS operator the ability to choose the appropriate CL during cruise, whether maximum range (.79), maximum endurance (1.35), or minimum fuel burn (.96).

Figure 3.2.1: Results from preliminary aerodynamic projections. (a) Preliminary drag polars for different aircraft mission segments. (b) Parametric analysis of lift and drag data. C_L corresponding to the maximum C_Lâ „C_D at constant Mach number maximizing the range. đ?‘Şđ?‘ł corresponding to the maximum đ?‘Şđ?‘ł đ?&#x;Ž.đ?&#x;“ đ?‘Şđ?‘Ť maximizes the range at a constant altitude. The parameter đ?‘Şđ?‘ł â „đ?‘Şđ?‘Ť đ?&#x;‘â „đ?&#x;? maximizes the SAR of the aircraft and was selected based on the recommendation given in ESDU 73019 as a measure of merit, defining a design region for the cruise đ?‘Şđ?‘ł of the aircraft.

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3.3) Constraint Diagram The aircraft is limited by how much thrust is required with respect to the weight of the aircraft, and how much weight can be supported by the wings. The goal is to obtain a configuration with the greatest wing loading possible so that the wing structural weight is minimized, while achieving as low a thrust loading as possible. To find the boundaries that limit how high the wing loading and how low the thrust loading can be pushed, the sizing of the aircraft was generated based on the performance requirement in MIL-STD using the methods presented by Roskam. Next the wing loading and thrust-to-weight ratio constraints were determined by solving performance boundary equations. Given takeoff constraints provided by RFP, the landing field length was assumed to be 7000 ft. The maximum lift coefficient was assumed to be 2.7 using trailing edge, single-slotted fowler flaps, according to the methods outlined in Fundamentals of Aircraft and Airship Design

by Nicolai. The Thrust and Wing loading limitations of each performance component are plotted below to display the allowable range of loadings (un-shaded area of the plot).

Figure 3.3.1: Constraint Diagram

Figure 3.3.1 shows that the thrust-to-weight ratio and the wing loading are limited by the requirement for stall (100 knots) and take-off CL max (2.2), with a resulting optimal thrust to weight ratio of 0.167 lb/lb and a wing loading of 81.55 lb/ft2.

4) Configuration Description 4.1) Overall As a starting point for any aircraft design, the decision between a traditional tube and wing versus a flying wing/blended wing-body configuration must be made. The traditional tube and wing structure generally yields better aerodynamic performance when compared to a flying wing due to the large thickness ratio required to make a flying wing capable of containing its payload. The flying wing configuration is more structurally efficient than the tube and wing configuration because it eliminates the need for joints to connect 17 | P a g e

the wing to the fuselage thus creating a localized stress concentration. An additional consideration that is specific to the design of the UAS is the required azimuth angles of the laser turret. The requirement for ¹120° of azimuth would require a flying wing configuration to have a leading edge sweep angle of 30° to avoid interference with the laser. Though the azimuth angle of the laser also affects the tube and wing configuration, the leading edge sweep is a function of the distance between the nose of the aircraft and the leading edge of the wing.

The above considerations led to the decision to utilize a tube and wing configuration. The minimum dimensions of the laser when paired with the root chord of the wing with an aspect ratio of 15, would require an excessive thickness ratio for the root chord airfoil to accommodate the laser and its associated cooling equipment. Additionally, in order to reduce the thickness ratio at the root, the taper ratio of the wing to minimize the thickness of the root chord would result in a very short chord length along the outer span reducing available fuel capacity necessary for the required loiter performance. With the tube and wing configuration, the wing thickness could be optimized for aerodynamic performance and not be constrained by the height requirements to house the laser. The laser itself would be placed inside the fuselage and drive the sizing of the cross section with the power plant should it be placed inside the fuselage. The tube and wing configuration also simplifies the control surfaces configuration. Unlike a tube and wing configuration, a flying wing does not have control surfaces for pitch and yaw separated from the wing, which requires control surfaces and high lift devices to take on additional control roles to control pitch and yaw of the aircraft. By having the control surfaces separated from the wing, the pitch and yaw axes would have control surfaces dedicated to them.

After selecting the tube and wing configuration, the last high level design decision was between a canard and a traditional empennage configuration. A canard configuration is advantageous from an aerodynamic perspective because canards do not create a downward load on the aircraft to pitch up like a traditional empennage does. However, a canard configuration is disadvantageous from a control and design perspective. A canard, unlike a traditional empennage, does not generate a self-correcting moment to bring the aircraft back to equilibrium and thus requires significantly more pilot workload which in this case is irrelevant.

The other issue with a canard configuration is that it has to be designed so that it stalls before the wing does so that pitch control authority of the aircraft can be maintained, this in turn limits the maximum lift that can be attained by the wing.

Despite the limit on useable lift that the canard imposes, a canard configuration was chosen for its aerodynamic benefits, as well as the resulting rearward shift of the wing when compared to a traditional empennage. This allowed for greater flexibility in designing the wing planform so that it did not interfere with the azimuth requirements of the laser. In the event of a catastrophic failure of the flight control system, the aircraft would become unstable in the pitch axis, potentially causing a mid-air breakup of the aircraft if the gloading of the aircraft exceeds the design limit. However the probability of this occurring is remote due to the level of redundancy that is in place in the flight control system.

4.2) Wing The wing planform has an area of 2580.375 sqft and a 196.6 ft span, resulting in an aspect ratio of 15.

In order to optimize the wing for loiter performance a spreadsheet was created with wing loading and wing coefficient of lift as independent variables. Using both the wing loading and the coefficient of lift, the lift to drag ratio of the wing was calculated in conjuction with assumed constants across the trade study space such as aspect ratio, oswald efficiency factor, and zero-lift drag coefficient. The resulting lift to drag ratios were then compared to the mathematical maximum to determine the optimal loiter wing loading of 68 lbs/sqft.

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Table 4.2.1: Loiter Wing Loading Optimization Spreadsheet. Cells with “False” represent lift to drag ratios inconsistent with theory

0.7 0.8 0.9 1 1.1 1.2 1.3 1.4

65 False 24.31032 21.53402 19.09657 16.97314 15.13049 13.53349 12.14864

66 False 24.68433 21.86532 19.39037 17.23427 15.36326 13.7417 12.33554

67 False 25.05833 22.19661 19.68416 17.49539 15.59604 13.9499 12.52244

68 False 25.43234 22.5279 19.97795 17.75652 15.82882 14.15811 12.70935

69 False False 22.85919 20.27175 18.01764 16.06159 14.36632 12.89625

70 False False 23.19049 20.56554 18.27877 16.29437 14.57453 13.08315

The weight estimation for the beginning of the loiter phase was then divided by the optimal loiter wing loading to generate the design wing area of 2580.375 sqft. From this design wing area, the takeoff wing loading and wingspan was calcualted to be 78.7 lbs/sqft and 196.6 ft respectively. This large wingspan was necessary to achieve the design aspect ratio and improve loiter performance. The wingspan is smaller than the Air Force’s largest tactical airlift aircraft the Lockheed C-5 Galaxy with a wing span of 222.9 ft [4]. This allows the UAS to utilize hangars already in place at any base that houses an Air Force Airlift Wing equipped C-5 Galaxies. If the UAS is deployed to a location without the hangars for the C-5 Galaxy, the appropriate hangars could be constructed using already existing designs and would expand capabilities of the airbase that the UAS is deployed to. A taper ratio of 0.4 was selected to be in upper-mid range of similar military aircraft as well as to limit wingtip stall to provide effective aileron response throughout the flight regime without sacrificing oswald efficiency. The quarter chord of the wing has a 4.9° sweep angle, in addition with the taper ratio, the root and tip chords are 18.76 ft and 7.5 ft respectively. The wing employs a NACA 64-814 airfoil throughout the span without any geometric twist.

Figure 4.2.1: Wing Planform and Dimensions in feet

The wing is supported by 3 spars placed at 10%, 50%, and 75% of the chord and by ribs along the span placed 24 in apart perpindicular to the trailing edge, common among military bombers, patrol, and transport aircraft [5]. The wing contains an excess of space to house the 17,413 gallons of fuel that would be stored within the wing. The wing ribs have holes along the chord to facilitate the flow of fuel through the wing tanks while minimizing loss of structural strength.

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To provide additional lift during take-off and landing, the wing is equipped with fowler flaps. Fowler flaps were chosen for their higher performance when compared to plain, split and single slotted flaps and their reduced mechanical complexity when compared to double and triple slotted flaps that offer similar

Figure 4.2.2: Wing Structure and Locations of Fuel Tank (Green) and Flaps (Blue)

performance. The flaps extend from 11.9% to 71.2% of the span and begin slightly behind the rear spar to allow space for the extenstion and deflection mechanisms to be mounted to the rear spar. The ailerons extend from the termination of the flaps to the wingtip. The ailerons were sized to be in the upper range of aileron areas among similar aircraft to maintain aileron effectiveness despite the large wingspan which would slow the roll response of the aircraft. During take-off and landing spoilers will be used for roll control in addition to the ailerons until the aircraft has achieved enough speed that the ailerons have enough control power to control the roll of the aircraft by themselves.

4.2.1) Landing Gear The main gear are 12 ft in length extending from the wing to the ground, have a strut diameter of 3.6 in and are manufactured with 5454 Aluminum (length of strut is 124 in). The length was chosen in order to provide aft fuselage clearance in case of over rotation. The material and diameter of the struts were determined using stress analysis of the loading that would be applied to struts with a factor of safety of 2. The main landing gear are positioned 30 ft from the center of the fuselage in order to provide 10째 wing tip clearance for landing and takeoff. The landing gear struts begin 60 in forward of the wing trailing edge between the aft and middle spar. When stowed, the landing gear will retract at an angle towards the leading edge of the wing.

There will be an interruption in the middle spar in order to stow the tires, which will sit between the forward and aft spars. Each main gear has four tires with a 40 in diameter and a 12 in width with a tire pressure of 180 psi.

The nose gear is positioned 25.7 ft aft of the nose and retracts aft into the fuselage. It has a length of 11 ft 4 in and has a strut diameter of 2.5 in. The nose gear has two tires with a diameter of 30 in, a width of 8 in, and tire pressure at 180 psi.

4.2.2) Aspect Ratio Justification As demonstrated in aircraft that are characterized by high altitude and long endurance such as the Global Hawk and AeroVironment Helios, high aspect ratios are used to create the most efficient aircraft. But such high aspect ratios give rise to the phenomenon of flutter. Flutter as defined by Nicolai is the dynamic instability of an elastic structure in an airstream, which occurs when the phasing between motion and aerodynamic loading extract an amount of energy from the airstream is equal to the energy dissipated within the wing from damping. As discussed in section 4.2, the UAS is designed for an aspect ratio of 15, which is not quite the aspect ratio of the Global Hawk (25.3) [6] but is greater than similar class airplanes of the B787 20 | P a g e

(9.59) [7] and B757 (7.8) [8]. To calculate the flutter limits of the aircraft, a conceptual flutter analysis MATLAB program by Vivek Mukhopadhyay [9] was used. The program uses basic geometric, structural, and mission inputs to determine the flutter region. The basic geometric inputs include:  đ?‘?đ?‘&#x; – 18.76 ft. Sec 4.2  đ?‘?đ?‘Ą –7.5 ft. Sec 4.2  S – 98.3 ft. Sec 4.2  ∆ - 4.9 deg Sec 4.2  đ??¸đ??´đ?‘™đ?‘œđ?‘? – 35% of chord The basic structural inputs include:  

GJ – root: 1.76015đ??¸10 đ?‘™đ?‘?â „đ?‘“đ?‘Ą 2 Matweb [10] đ??źđ?‘?đ?‘–đ?‘Ąđ?‘?â„Ž - 1.0đ??¸6 đ?‘™đ?‘? ∗ đ?‘“đ?‘Ą 2 Sec 11.1

 đ?‘Šđ?‘’đ?‘Ľ – 25000 lb Sec 11. 1 and mission inputs including:  Mach number (.56 for loiter and .722 for cruise Sec 3.1)  Cruise altitude (40,000 ft Sec 3.1) The program begins it analysis by calculating the Regier and Flutter numbers. The Regier number is defined by Vivek as the ratio of elastic force over aerodynamic force at sea level and calculated as đ?‘… = đ?‘‰đ?‘… â „đ?‘Ž0 Where đ?‘Ž0 is speed of sound at sea level and đ?‘‰đ?‘… is the Regier Surface Velocity Index, defined as đ?‘‰đ?‘… = 0.5đ?‘?75 đ?œ”đ?›ź √đ?œ‡0 Where đ?‘?75 is the chord at 75% semi span, đ?œ‡0 is mass ratio at sea level, and đ?œ”đ?›ź is the torsional frequency which is defined as đ?œ”đ?›ź =

đ??žđ?‘Ž đ??şđ??˝đ?‘&#x;đ?‘œđ?‘œđ?‘Ą √ đ??ż đ??ź60 â „đ?‘”

Where đ??žđ?‘Ž is a factor relating the GJ ratio, L is the effective spar length, đ??şđ??˝đ?‘&#x;đ?‘œđ?‘œđ?‘Ą is the torsional stiffness at the root, đ??ź60 is the running pitching moment at 60% semi span, and g is the gravitational constant.

The Flutter number is defined by Vivek as equivalent air speed at sea level divided by the surface velocity index or the Mach number divided by the Regier number đ??š = đ?‘€ â „đ?‘… With the two defining parameters calculated, the program varies the numbers with altitude, producing boundary plots as seen in Figure 4.2.3: Flutter and Regier number boundaries as a function of Mach number for low sweep angles.

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Figure 4.2.3: Flutter and Regier number boundaries as a function of Mach number for low sweep angles

And combining the two boundary plots, a summary boundary plot is constructed with Mach number and dynamic pressure used for the x and y axis accordingly as seen in Figure 4.2.3: Flutter and Regier number boundaries as a function of Mach number for low sweep angles As denoted in Figure 4.2.4 by the two x, the loiter and cruise mission segments respectively are below the flutter dynamic pressure boundary of 40,000 ft. Since both segments happen at 40,000ft, the UAS is able to fly at desired cruise altitude. Vivek notes that for safety an aircraft should have a 44% margin between the flutter and dynamic pressure boundaries. Since the dynamic pressure boundary for cruise and loiter flight segments is lower that the flutter limit set by the material by a margin greater than 44%, the UAS will not have any risk of encountering flutter as primary result of failure. As noted by Sarhaddi [11], a main symptom of structural failure is the reduction of the torsional frequency.

The addition of a middle spar as discussed in section 4.2), which was not accounted for in the MATLAB program, will further strengthen the stiffness of the wing and increase the flutter boundary. A solution, as noted by Sarhaddi, is to reduce the flexibility of the root of the wing by strengthening the structure that holds the wing in place. With the choice of Aluminum 2024-T3 as the material of spar as discussed in section 9.1), the wing will get the additional stiffness without suffering too great a weight penalty.

The discussion of the structure of the wing box is discussed in section 4.2). So as result, the primary driver in reducing the torsional frequency is the pitch moment of inertia, Figure 4.2.4: Summary table with flutter and dynamic which is discussed in section 11.3). A higher pressure boundaries as a function of Mach number. The pitching moment reduces the torsional two x show the dynamic pressure of the aircraft in it’s loiter and cruise conditions respectively at an altitude frequency of a given wing size and material, leading to a greater probability of flutter of 40,000 ft.

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occurring via a structuralAnd combining the two boundary plots, a summary boundary plot is constructed with Mach number and dynamic pressure used for the x and y axis accordingly as seen in Figure 4.2.3: Flutter and Regier number boundaries as a function of Mach number for low sweep angles As denoted in Figure 4.2.4 by the two x, the loiter and cruise mission segments respectively are below the flutter dynamic pressure boundary of 40,000 ft. Since both segments happen at 40,000ft, the UAS is able to fly at desired cruise altitude. Vivek notes that for safety an aircraft should have a 44% margin between the flutter and dynamic pressure boundaries. Since the dynamic pressure boundary for cruise and loiter flight segments is lower that the flutter limit set by the material by a margin greater than 44%, the UAS will not have any risk of encountering flutter as primary result of failure. As noted by Sarhaddi [11], a main symptom of structural failure is the reduction of the torsional frequency.

The addition of a middle spar as discussed in section 4.2), which was not accounted for in the MATLAB program, will further strengthen the stiffness of the wing and increase the flutter boundary. A solution, as noted by Sarhaddi, is to reduce the flexibility of the root of the wing by strengthening the structure that holds the wing in place. With the choice of Aluminum 2024-T3 as the material of spar as discussed in section 9.1), the wing will get the additional stiffness without suffering too great a weight penalty. The discussion of the structure of the wing box is discussed in section 4.2). So as result, the primary driver in reducing the torsional frequency is the pitch moment of inertia, which is discussed in section 11.3). A higher pitching moment reduces the torsional frequency of a given wing size and material, leading to a greater probability of flutter occurring via a structuralmaterial, leading to a greater probability of flutter occurring via a structural failure. To reduce the pitching moment, the wing and tails are sized in order to reduce the pitching moment by optimizing the distance between the respective aerodynamic center and the aircraft center of gravity, which is discussed in section 11.3).

4.3) Fuselage The diameter of the fuselage was driven by the laser diode cross section of an 8.2 ft by 8.2 ft square.

The laser diode had to fit within the fuselage ribs with room to fit the cooling system and supporting structure.

The length of the aircraft was set at 170 ft to provide a long enough moment arm for the pitch and yaw control surfaces, so that the surfaces were not excessive in proportion to the rest of the aircraft. In the process of determining the fuselage length, models of the proposed fuselage were created with the wing placed near the approximate location of the center of gravity and the empennage sized to the wing and fuselage combination.

A 150 ft long fuselage was considered initially, however the empennage was excessively large in comparison to the rest of the aircraft. A 180 ft long fuselage was then considered and determined to allow for sufficient empennage size, but it would add unnecessary weight to the configuration because the extra space it provided was not needed. A 170 ft fuselage was settled on as the appropriate balance between weight and empennage size. Due to this large diameter fuselage cross section and the fuselage length, the aircraft has an abundance of empty space within the fuselage that could be utilized for future capability upgrades at the expense of loiter time.

Figure 4.3.1: Fuselage Dimensions

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The fuselage has two types of cross section over its length, circular and double ellipsoid. At the point the turret is mounted the fuselage has a 6 ft. circular cross section which then merges into a double ellipse with a bottom half with a 13 ft. diameter circle and the top half contains an ellipse with a major axis radius of 8.5 ft.

The double ellipsoid section contains a radome to provide capacity for the communications and control avionics to function to their full capacity while being sheltered from aerodynamic forces and separated from the 1 MW infrared laser beam that traverses the fuselage. The radome section transfers into the main circular section of the fuselage, carrying on the 13 ft. diameter of the lower portion of the radome section. At 150 ft. fuselage station the fuselage tapers off from 13 ft. to 7 ft. in diameter to serve as a nozzle for the single centerline mounted engine.

The fuselage’s structure is made up of ribs that are spaced 24” apart and are 4” deep as specified in Roskam. The ribs follow the external shape of the fuselage except for the section that supports the radome.

Figure 4.3.2: Fuselage Structure

The radome ribs have a flat top on which large avionics systems can be mounted easily to the fuselage structure. Longerons are spaced around the circumference of the airframe with three sets of longerons spaced 36” apart centered at 3 o’clock and 9 o’clock with additional longerons at 12 and 6 o’clock. Additional longitudinal strength comes from the carbon composite skin of the aircraft. The wing, and empennage are attached to the fuselage through ribs that have a rectangular outcropping from the circular cross section that attaches to wing or empennage ribs.

4.4) Canard and Vertical Tail The canards each have an area of 182.5 sqft and a span of 21.35 ft for an aspect ratio of 2.5. The quarter chord sweep of the canards is 11.5°. Because canards are rarely employed on aircraft, typical values for horizontal tails were applied towards the design of the canard. To control the pitch of the aircraft, elevators are attached at the 75% chord location and extend from the 15% span location to the tip of the canard. The canards are mounted slightly above the centerline and significantly ahead of the wing to prevent the canard’s wake from interfering with the wing’s flow field during level flight. The vertical tail is mounted in the rear of the aircraft, so that an overly complex control scheme for the lateral axis of the aircraft is unnecessary. The vertical tail has an area of 600 sqft and a span of 30 ft. for an aspect ratio of 1.5 which is within the range of

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Figure 4.4.1: Vertical Tail and Canard Structure

aspect ratios for vertical tails of similar aircraft. The quarter chord sweep of the vertical tail is 17.75° is within the range of similar aircraft as well. For yaw control, the vertical tail has a rudder mounted to the 75% chord location spanning from 15% of the span to the tip of the vertical tail for a rudder area of 122.8 sqft.

The ribs of the vertical tail are parallel to the central axis of the fuselage from 0% span to 15% span, where the rudders begin. From 15% span to the tip of the vertical tail, the ribs are perpendicular to the leading edge. Parallel ribs are spaced 24” apart and the first 2 to 4 ribs that are perpendicular to the leading edge intersect the last rib parallel to the fuselage. The ribs of the canards are all parallel to the central axis of the fuselage and also have 24” between each rib. The 24” spacing reflects the wing rib spacing criterion referenced earlier and applies to the vertical tail and the canards since they are merely small symmetric wings that provide a stabilizing force when the control surface is trimmed. The vertical tail and the canards each contain 2 spars, one at 15% chord location and the other at the 75% chord location to provide a hinge line for

Figure 4.4.1: Vertical Tail and Canard Planforms

the respective control surface.

4.5) Propulsion A high bypass turbofan was chosen to power the UAS for a multitude of reasons, first among them is efficiency. Because the aircraft is required to loiter for a minimum of 24 hours, an efficient engine is needed to reduce the amount of fuel that needs to be carried and thus the structure required to support that weight. The only types of engine that are efficient enough to provide the level of endurance that is required by the RFP are turboprops and turbofans which accelerate a large mass of air less than a low bypass turbofan or turbojet which accelerate a smaller mass of air to a faster speed. Another reason that a high bypass turbofan was selected was its performance at altitude. Air density decreases with altitude and thus decreases the amount of thrust that can be generated. Turbofans generally tend to have better high altitude performance than turboprops because of the loss of density. A more detailed comparison between the turbofan and turboprop engines is presented in section.

A General Electric CF6-80E1A4, a two spool, “dual Rotor, axial flow, annular combustion chamber, high bypass turbofan” [12] was used for this purpose. The CF6-80E1A4 has the following dimensions.

Dimensions Overall Length 168.41 [inches] 25 | P a g e

Overall Width Overall Height Overall Weight (Dry)

114.13 [inches] 113.13 [inches] 11,225 [lb]

As the CF6-80E1A4 engine generates enough thrust to propel the UAS with the use of solely one engine, only one engine will be implemented in order to reduce weight, which will be imbedded in the aft of the fuselage in order to reduce drag, as shown below.

5) Aerodynamics 5.1) Airfoil Selection Two main elements dictated the method of the selection of airfoil profiles. First, in order to meet the lift coefficient requirement for desired cruise and loiter performance, the profile had to achieve a lift coefficient of 1.1. Yet for the aircraft to maintain level flight, the airfoil profile had to achieve a lift coefficient of 1.1 with 3 degrees in a 2-D analysis. Secondly, the airfoil geometry must have sufficient volume to house the wing structure and fuel volume. The limits for the thickness-to-chord ratio were set to 14% for the root chord and 9% for the tip chord based on drag divergence Mach number considerations as shown in Sect X. In order to obtain a baseline airfoil and considerations of the loiter and cruise phase, a study of 10 high lift airfoils, available on the University of Illinois Urbana – Champaign’s web portal and DesignFoil software’s NACA airfoil generator. The airfoils were analyzed using the XFOIL software on the merit of the lift coefficient generated at low angles of attack at the Reynolds numbers of the cruise and loiter flight conditions.

The NACA 64814 airfoil was selected as profile at both the root and the tip. CFD analysis using Star CCM+ software was performed to verify the lift coefficient produced at low angles of attack for the chosen airfoil profile. The analysis was performed at an altitude of 45,000’ and a Mach number of .64 for cruise and .46 for loiter, which included a 2 deg of wing incidence. The mesh used was generated using a surface remesher and an advanced layer (volume) mesh models. Turbulence was modeled by an SST K-Omega turbulence model.

When 2D analysis was performed, free stream and symmetry planes were used to create a 2D infinite wing.

Figure 1 and Figure 2 present the results of the CFD analysis performed for the selected airfoil. The airfoil profile for the horizontal tail was chosen to be a symmetric to help with the stability and control of the aircraft.

With a desired thickness to chord ratio of 14% for the horizontal tail, the NACA 64-014 profile was selected.

5.2) High Lift Devices Since the design called for a large and heavy aircraft that needs to be able to take off and land quickly on a 7000 ft runway, high lift systems were a key design feature. To provide additional lift during take-off and landing, the wing is equipped with fowler flaps. Fowler flaps were chosen for their higher performance when compared to plain, split and single slotted flaps and their reduced mechanical complexity when compared to double and 26 | P a g e Figure 5.2.1: Maximum Lift Coefficient vs. Flap Deflection at for several flap chord to wing chord ratios

triple slotted flaps that offer similar performance. The flaps extend from 11.9% to 71.2% of the span (thus taking up 59 % of total span) and begin slightly behind the rear spar to allow space for the extension and deflection mechanisms to be mounted to the rear spar. Increasing the flap chord with respect to the wing results in increased lift, but the flap length is limited by the position of the rear spar to a maximum flap chord to wing chord ratio of 0.25. Thus the resulting total flap platform area is 355 ft2 (14% of the total wing area). The maximum deflection of the flaps to be used during takeoff is δ=12°, resulting in a takeoff CL of 2.4, while during landing the maximum deflection is δ=40° resulting in a CLmax of 2.7. These design points are shown in figure 5.2.1 which also displays the effect on CLmax due to increasing the deflection angle and the percent of total chord. The maximum deflection angles to be employed were determined from the specifications detailed in UMD lectures. Since the flaps provided sufficient lift to meet design constraints, leading edge slats were not included as they would require additional complexity cost.

The ailerons extend from the termination of the flaps to the end of the half span. The ailerons were sized to be in the upper range of aileron areas among similar aircraft to maintain aileron effectiveness despite the large wingspan which would slow the roll response of the aircraft. During take-off and landing spoilers will be used for roll control in addition to the ailerons until the aircraft has achieved enough speed that the ailerons have enough control power to control the roll of the aircraft by themselves.

5.3) Detailed Drag Polars To generate a more accurate estimate of the lift and drag forces experienced on the aircraft, a more in depth analysis of the aerodynamics was performed using methods described by Roskam [13]. The methodology used to determine the cruise drag polars accounts for compressibility effects by using the corrections displayed ESDU Transonic Aerodynamic Items 6407, 71019, 79004, and 83017 [3]. The low speed drag polars used the methodology as presented by Torenbeek [14]. Figure 5.3.1 presents the results of the detailed drag analysis using 5th order drag polar equations, which will be used in the Performance section to

Figure 5.3.2: Drag Breakdown at Cruise. The 12% of drag generated by the fuselage shows the successfully implantation of the laser with respect to the performance of the aircraft

27 | P a g e Figure 5.3.1: Drag Polars at key mission segments. Cruise and Loiter points depict optimal flight performance. Takeoff and Landing points indicate takeoff and landing at a CLmax of 2.2 and 2.7 respectively

verify performance requirements. The drag breakdown at cruise is presented in Figure 5.3.2.

5.4) Computational Fluid Dynamics Computational fluid dynamics software was used throughout the design phase of the aircraft. It was used to assist in airfoil selection and provide validation of expected lift and drag coefficients. Next, CFD was applied to a study to determine whether a configuration incorporating a canard stabilizer would be preferable or not. CFD was then used to analyze the final aircraft design for expected lift-to-drag ratio, and evaluate flow field parameters such as pressure profiles and turbulence distributions.

5.4.1) Airfoil Validation A number of airfoil profiles were selected as candidates for good subsonic performance at high altitude. Expected lift and drag coefficients were determined using airfoil design software such as XFoil.

These coefficients could then be compared to a CFD 2-D analysis for consistency. For the 2-D airfoil analysis, expected cruise flight conditions of 40000 ft altitude and Mach 0.5 airspeed were used, with angles of attack typically at 0 or 10 degrees. The coefficient curves are expected to be linear in the range between 0 and 10 degrees.

A simple turbulent flow model, k-omega, was used for all CFD runs. It has slightly better accuracy than k-epsilon, and does not have a large computation time penalty. For accurate modeling of the boundary layer with turbulent flow enabled, good inflation layers must be used such that the Law of the Wall can be captured in the simulation. The Law of the Wall states that turbulent flow average velocity is logarithmically related to the distance of a flow region to the wall. Wall y+ is the dimensionless distance to the wall, and for values between 5 and 30, there is a transition to a linear relation between average velocity and wall distance.

Thus it is best to design a CFD mesh which can model the boundary layer of the flow field down to the viscous sublayer, where there is a linear relation between wall distance and average flow velocity. This occurs at y+ values below 5, and ideally should be 1. An example mesh from the analysis for one particular airfoil with wall y+ plot is below.

Figure 5.4.1: Example CFD Cell Configuration

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5.4.2) Canard Study CFD was used to determine possible performance improvement from using a canard configuration rather than a conventional aircraft with horizontal stabilizer mounted at the rear of the aircraft. For the study, a 3-D model of a fuselage section, forward canard, and wing were used. The main wing had a chord length of 6 meters, and the canard had a chord length of 4 meters.

A canard was placed forward of the wing by either 3 or 4 times the chord length of the main wing. It was also raised or lowered relative to the chord line by either -1/12, +1/12, or +1/4 times the chord length above or below the main wing. Thus, cases were evaluated for combinations of the above (8 in total, with one without a canard) in search of improvement to the lift-to-drag ratio of the main wing, and to find which cases have the canard wake interfere with the main wing. Canard incidence angle and elevator deflection were also considered. Lift-to-drag ratio results are tabulated below: Table 5.4.1: Summary of Canard Positions Studied in CFD

Canard Forward Distance 18 m 18 m 24 m 24 m 24 m

Canard Elevation +1.5 m +1.5 m +1.5 m +1.5 m -0.5 m

Aircraft Angle of Attack 3 degrees 3 degrees 3 degrees 3 degrees 3 degrees

24 m

-0.5 m

3 degrees

24 m

-0.5 m

3 degrees

Canard Incidence Angle 0 degrees 10 degrees 0 degrees 10 degrees 0 deg / +15 deg elevators 0 deg / -15 deg elevators 3 degrees

Wing Incidence Angle 2 degrees 2 degrees 2 degrees 2 degrees 2 degrees

Lift to Drag Ratio

2 degrees

24.0

2 degrees

23.8

24.6 23.1 24.8 23.5 30.1

Sample CFD images are below, comparing the case with L/D = 30.1 with the no canard case.

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Figure 5.4.2: Canard Below Wing CFD Result

Figure 5.4.3: Wing CFD Analysis

With no canard present, the lift to drag ratio was 26.7. The only significant increase to L/D was found when the canard was placed -0.5 m below the main wing chord line, and the elevators were deflected up 15 degrees. The study did not suggest a sustained benefit to L/D from using a canard, but rather a consistent reduction in L/D by about 5 to 10%. Following further aerodynamic analysis without CFD, the final configuration was selected to not have a forward canard.

5.4.3) Analysis of Final Aircraft Configuration CFD was used to validate the expected aircraft performance calculated during the design phase. The following flight conditions were used: Mach = 0.515, Altitude = 40000 feet. Atmospheric properties used in the simulation were obtained based on the 1976 United States Standard Atmosphere model.

Turbulence modeling (k-omega) and the energy equation were enabled, and the air density was modeled with the ideal gas law. A half-cylinder flow field was constructed around one-half of the aircraft model so that symmetry could be used. The fluid volume was scaled to be roughly 10 times the length of the aircraft to keep the inlet and outlet boundaries to the far-field. A mesh of about 6 million cells was used, with a simple 10 element thick inflation layer around the wing, stabilizers, and fuselage. Due to computer constraints, inflation could not be added to the ailerons or elevators.

The final aircraft design was simulated in a cruise configuration with angle of attack set to 3 degrees.

The table below contains body force results, totals of lift and drag forces, and the overall L/D ratio.

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Table 5.4.2: Summary of CFD Results

Body Forces [lbf] Wing Aileron Fuselage H-Stab V-Stab Total Force [lbf]

Lift 71269 595 5363 -526 377 148792

Wing L/D Ratio Total L/D Ratio

37.4 35.8

Drag 1903 48 896 -233 -85 4161

The 3D CFD simulations of Ancile are presented below at cruise conditions. For numerical calculation simplicity, the 3D CAD model was sliced longitudinally along the center line and a symmetry plane was added in place of one half of the model. When applicable, the full 3D CAD model was used to generate simulations showing the flow around the whole aircraft. Simulations show the flow moving around the turret on the nose with little penalty, allowing subsonic and transonic predictions to be used without considering the turret. As shown below, the effects from the canard tip vortices go above the wing and the inlet, avoiding harmful influence to the performance of the wing and engine. The showing of minimum drag in 3D CFD simulations gives SCAD the ability to stand behind predicted lift-to-drag ratios.

Figure 5.4.4: Sample CFD Results

6) Propulsion 6.1) Turbofan versus Turboprop During the preliminary design stages two forms of propulsion were considered for use on the UAS: high-bypass turbofans and turboprops. Both types of propulsion were feasible for the design flight speed, 31 | P a g e

however for a turboprop engine special propellers may have to be implemented as portions of the propeller blade could go supersonic. The Russian Tu-95 [13] has propeller blades that allow the aircraft to travel at approximately 500 knots, well in excess of what the aircraft would be flying at to complete its mission. To aid in the decision, preliminary weight estimates were run for both turbofan and turboprop configurations based on readily available specific fuel consumption figures for existing engines. One initial issue with turboprops was that the most powerful turboprop was produced by the Russian Federation for its Tu-95 bomber and would not be feasible to implement on an American aircraft for strategic purposes. The most powerful turboprop produced by a NATO member was significantly less powerful and based on calculations to compare the power produced by turboprops to the thrust produced by jet engines, it was determined that an increase in the number engines would be necessary to achieve the same propulsive force, even with the more powerful Russian engine. However, because the Russians had built a turboprop that would have fit the needs of the UAS had it been produced by a NATO member and technology had improved since the introduction of that engine, turboprops were not eliminated at that stage. Upon completing the weight estimates for turbofans and turboprops, the difference between the TOGWs was in favor of the use of turbofans.

The main factor driving the difference was the amount of fuel required by the different configurations.

The turboprop configuration used a larger amount of fuel, but had a lower structural weight than the turboprop.

The reduction in structural weight was due to the turboprop configuration being optimal at a lower wing loading and lift coefficient, resulting in a much larger wing than the turbo jet configuration. Over the lifetime of the aircraft the increase in fuel costs would greatly exceed the initial savings in material costs. Lower wing loading also tend to be less efficient in utilizing lift and add additional surface area which affects the amount of drag, an additional factor in increasing the fuel load when compared to the turbofan. A turbofan configuration was chosen over the turboprop to reduce mechanical complexity, cost, and because it resulted in a lower TOGW.

6.2) Engine Selection and Characteristics Due to the transonic nature of flight, the UAS would optimally be propelled with the use of a high bypass turbofan with a low SFC, high net thrust, and moderate specific thrust. As such commercial grade turbofan engines were considered. Of these engines, the General Electric CF6-80E1A4 was chosen due to its similarity to the engine used to propel the C-5 Galaxy [14], low SFC, high static thrust rating, and most importantly, due the fact that its diameter more closely matches to the fuselage diameter necessary to house the laser. The GE CF6-80E1A4 is a two spool, “dual Rotor, axial flow, annular combustion chamber, high bypass turbofan� with a 14-stage High Pressure Compressor (HPC) run by a two-stage High Pressure Turbine (HPT) and a three-stage Low Pressure Compressor (LPC), with integrated one-stage fan, run with a five-stage Low Pressure Turbine (LPT). The engine is then controlled using a “Dual Channel Full Authority Digital Electronic Engine Control Unit (FADEC)� [12] and has a design bypass ratio of 5.1, yielding the following published static thrust and SFC characteristics at sea level.

Table 6.2.1: Engine Static Characteristics

Takeoff Thrust Maximum Continuous Thrust Inlet Area Pressure Ratio SFC

Static Characteristics 66,870 [lbf] at Sea Level 60,400 [lbf] 70.846 [ft2] 32.4-34.8 0.332-0.34

The thrust characteristics can be found analytically using the density ratio, đ?œŽ, where đ?œŽđ?‘Žđ?‘™đ?‘Ą =

đ?œŒđ?‘Žđ?‘™đ?‘Ą đ?œŒđ?‘†đ??ż

, for

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both cruise and high elevation takeoff conditions. Where, đ?‘‡đ?‘?đ?‘&#x;đ?‘˘đ?‘–đ?‘ đ?‘’ = đ?œŽđ?‘?đ?‘&#x;đ?‘˘đ?‘–đ?‘ đ?‘’ ∗ đ?‘‡đ?‘†đ??ż = đ?‘‡â„Žđ?‘’đ?‘Ą = đ?œŽâ„Žđ?‘’đ?‘Ą ∗ đ?‘‡đ?‘†đ??ż =

đ?œŒđ?‘?đ?‘&#x;đ?‘˘đ?‘–đ?‘ đ?‘’ ∗ đ?‘‡đ?‘†đ??ż đ?œŒđ?‘†đ??ż

đ?œŒâ„Žđ?‘’đ?‘Ą ∗ đ?‘‡đ?‘†đ??ż đ?œŒđ?‘†đ??ż

where the subscripts SL and het correspond to sea level and high elevation takeoff respectively.

Thus the following thrust characteristic where generated for cruise as well as the takeoff conditions at an elevation of 6000 ft where it is unlikely for the UAS to be employed from a base with an elevation higher than 6000 ft.

Table 6.2.2: Engine Characteristics at High Elevation

Thrust [lbf] Cruise/Loiter (40,000 ft) 14,869 High Elevation Takeoff (6000 ft) 55,897 As the thrust required for cruise and high elevation takeoff is 13,315 and 44,222 lbf respectively, as well as 35,096 lbf to takeoff at sea level, the CF6-80E1A4 provides adequate thrust for the UAS to undertake its mission.

The CF6-80E1A4 was then modeled with the use of the program GasTurb, in order to give a better estimation of its performance during different sections of its mission. GasTurb generated the following results.

Figure 6.2.1: Carpet Plot of SFC as a Function of Thrust It is important to note, that due to the scarcity of published engine data, aspects of CF6-80E1A4 engine were estimated during its modeling in GasTurb. This, however, generated inherent uncertainties in its calculated values.

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Table 6.2.3: Summary of Engine Performance

Thrust Requirement [lbf]

Thrust Thrust SFC Available Available (GasTurb) (GasTurb) (Analytical) [lbf] [lbf] Sea Level 35,096 75,190.36 66,870 0.3131 6000 ft 44,222 62,025.56 55,897 0.3232 Cruise 13,315 13,998.58 14,869 0.5645 Loiter 8,006 13,585.44 14,869 0.5323 The thrust characteristics found both analytically and via GasTurb seem relatively consistent, especially during cruise and loiter. The General Electric CF6-80E1A4 engine is thus the optimal engine for the UAS as its approximately 14,000 lbf of cruising thrust allows it to overcome the UAS’s approximate 13,315 lbf of drag with enough of a buffer to account for weather and other turbulent conditions, while also producing enough lift to takeoff at any airstrip that the UAS is likely to be deployed from.

6.3) Engine Inlet Area The required inlet area is determined based on the engine mass flow, and the ratio of flow streamtube areas at various stages of flight to the inlet area. Subsonic inlets are typically designed with the following ratios:  Ground Operation – Ratio = Infinity  Cruise = Between 0.5 and 0.8.  Climb = 1.3 When the aircraft is on the ground moving at low speed, streamlines entering the inlet approach from all directions. This increases the streamtube area upstream of the inlet towards infinity as flow speed goes to zero.

When inlets are positioned along a surface (such as the fuselage), the boundary layer due to no-slip serves to restrict the flow of air into the inlet. A boundary layer splitter is used to provide separation between the inlet and the surface so that flow into the engine is restricted the least. Use of a boundary layer splitter allows the streamtube area ratio for the Climb phase of flight to be between 0.9 and 1.

The table below contains calculations for various phases of flight, from ground roll to cruise. The inlet area is chosen with iteration. The upstream streamtube area is calculated using the continuity equation, where w = ρVA. Mass flow is divided by aircraft velocity and the local air density. The area ratios listed below meet the criteria discussed above, with 0.91 at Cruise and 1.3 at Climb through 4000 feet.

Table 6.3.1: Engine Inlet Calculation

Input:

Calculation of Engine Inlet Area

Flight Stage Begin TO Takeoff Climb Cruise

Assumed Flight Conditions Velocity [ft/s] Altitude [ft] 10 250 350 500

0 0 4000 40000

Atmospheric Properties* Density P Static [lb/ft^3] [psi] 0.076474 0.076474 0.067917 0.018895

14.696 14.696 12.693 2.73

Inlet Area [ft^2]

Mach Number 0.009 0.224 0.318 0.517

55

**Engine Performance Thrust Mass Flow [lb] [lb/sec] 68568.5

11037

1900 1900 1696 474 34 | P a g e

Calculate Area Ratios A_∞ Ratio to Stage [ft^2] A_In Begin TO 2484.50 45.173 Takeoff 99.38 1.807 Climb 71.33 1.297 Cruise 50.17 0.912 * US Standard Atmosphere 1976

*Boundary Layer [inches] Stage

Laminar

Turbulent

Begin TO Takeoff Climb Cruise

8.92 0.564 0.501 0.715

38.31 12.7 12.1 14

** GE CF6-E8 Turbofan Performance

The boundary layer splitter must be at least about 1 inch wide. The laminar boundary layer width was calculated assuming 130 feet of fuselage surface upstream of the inlet.

7) Avionics 7.1) Introduction Avionics are included on almost every type of aircraft, including unmanned aerial vehicles (UAVs).

Most avionics are standard and perform satisfactorily, so that customized avionics are not necessary for the majority of aircraft. UAV’s, however, require some specific components that are not common to most aircraft, due to their remote control and operation. These components must be selected so that the aircraft can receive commands from a base station, act upon these commands, and function even if contact with the base station is lost. For example, a UAV must include satellite communications antennae and data links capable of transmitting and receiving data, to receive orders from a base of operations and to upload military intelligence, auxiliary data, and mission objective status updates. Robust and powerful onboard computer processing units are also required, since all commands should be routed through the main flight or mission computers. These computers are responsible for the smooth functioning of the aircraft, and must assume control of the aircraft if contact is lost with the home base of operations. Therefore, a UAV should have autonomous features, at least sufficient to recover the aircraft if contact with the UAV is lost.

Additionally, the lack of a pilot suggests that more sensor packages must be placed on a UAV to gather the same situational information available to a pilot.

These sensor packages must feed the information they gather into the central flight and mission computers on the aircraft, which serve as the “brain” for the aircraft. These computers control the flight pattern of the aircraft, and constantly check mission objectives to determine whether or not they have been fulfilled. A UAV should have a distributed and redundant computing system, Figure 7.1.1: The spaced distribution of the computer resources aboard an aircraft. This distribution prevents so that the entire aircraft is not lost if one or catastrophic failures if one or more computer units are destroyed or malfunction.

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more computers are damaged or malfunction (see Figure 7.1.1 below [15]).

Therefore, a UAV has some significant differences from a conventional, piloted aircraft. UAV’s must remain in constant contact with a home base of operations to receive commands and transmit location, status, and mission information, while piloted aircraft only require occasional contact. A UAV also needs software designed to control the aircraft and check mission status, since a pilot would perform those functions in a manned aircraft. Finally, pilots have a situational awareness of their airspace and surrounding terrain from cockpit instrumentation and transparencies, whereas a UAV requires sensor inputs for its flight and mission computers to receive and process situational data. These features all contribute to a different avionics load out for UAV’s as opposed to manned aircraft.

7.2) Requirements The avionics on this aircraft were selected primarily to satisfy the requirements of the anti-missile mission outlined in the RFP, and also to meet the basic requirements of a military aircraft. The avionics package includes basic and common components such as the mission control computer, flight control computer, radar, and SATCOM data link, as well as slightly more exotic components like the mission-specific infrared sensor and laser sensors used to pinpoint the enemy missile location.

Additionally, the UAV required certain avionics choices to be made to ensure the aircraft could operate remotely. For example, the aircraft required a robust SATCOM antenna, since the antenna would be the primary method of receiving orders and uploading the “intelligence, surveillance, and reconnaissance” (ISR) information. Likewise, the GPS/INS system is incredibly important, considering that the GPS/INS system accurately provides the location of the aircraft and is therefore critical to mission operation. Navigation of the UAV would be difficult without this GPS/INS system. Manned aircraft do not need to place such emphasis on the reliability of the SATCOM antenna or the accuracy and operation of the GPS/INS system, since the pilot(s) are capable of autonomously piloting the aircraft and completing mission objectives if contact with their home base is lost, and can navigate using landmarks, maps, or star positions as required. The inertial navigation system (INS) part of the GPS/INS serves as a means of tracking the position of the aircraft if contact with the positioning satellite is lost, but becomes inaccurate over long distances or time periods. This INS system is internally contained within the aircraft, and operates by double integrating linear and angular acceleration data over time to obtain the current aircraft position. This system helps combat “GPS spoofing,” or false GPS signals sent to the UAV’s GPS receiver to confuse the UAV as to its position. The required ISR systems are listed below, in Table 7.2.1.

Table 7.2.1: A basic list of ISR systems in this UAV.

System Name

Usage

1 KW Solid State Laser (tracking)

Tracks the target to assist with primary laser firings.

1 KW Solid State Laser (atm) Infrared sensor from RFP

Monitors atmospheric disturbances to assist with primary laser firings.

Tracks the target to assist with primary laser firings, and to determine whether the target was successfully destroyed.

Monitors aircraft altitude, which can be used for intelligence purposes.

Monitors aircraft’s surrounding environment; can also be used for ISR.

Radar Altimeter Radar

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RWR IFF

Can locate enemy aircraft or ground stations from their radar usage.

Can gather information about friend or foe positions and strength.

Besides this reliance on GPS/INS and other electronic navigation tools, this UAV must also gather intelligence about its surroundings through primarily electronic methods. These electronic methods include specific sensors, like the radar unit or the radar warning receiver (RWR). The information gathered using these methods is interpreted both by the UAV remote operators and the central flight control and mission computers.

This information is then used to make real-time tactical and strategic decisions for the UAV. For example, an incoming sortie of enemy fighters sensed using the RWR could cause the mission computer to enact an evasive flight pattern, and the remote operators could send commands to a nearby base to supplement the UAV’s escort with additional fighters. Therefore, the UAV’s electronic sensors provide actionable data and intelligence necessary for the successful operation of the aircraft. The major electronic sensors and devices are listed below Table 7.2.2.

Table 7.2.2: A basic list of electronic sensors and devices in this UAV.

Electronic Sensors & Devices Radar Altimeter

Radar RWR

IFF

GPS Radar Jammer KU SATCOM Antenna

ILS

Usage (besides ISR) Determines vertical position of the aircraft, used for flight path and navigation.

Detects obstacles or other aircraft, used for navigation.

Detects enemy radar usage, can be used to avoid those positions or for early warning of attacks.

Identifies friendly or enemy radar signatures to better understand the flight space environment.

Used to track position and navigate.

Used to inhibit enemy use of radar.

Used to upload and download commands and information from the aircraft to a satellite.

Assists in landing the aircraft.

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Figure 7.2.1: The front view of the fuselage contains the avionics placed towards the front of the aircraft, especially within the fuselage “bubble.”

The avionics listed in the two tables above (Table 7.2.1 and Table 7.2.2), as well as other avionics must be placed within the aircraft. The majority of the avionics were placed within the fuselage, with only some RWR sensor antennas and other sensors like the RFP-provided laser sensors being placed on the wings or other parts of the aircraft. The following figures contain the placement of the avionics components throughout the fuselage, displaying a front view, rear view, and overall view.

Figure 7.2.2: The rear view of the fuselage contains the avionics placed towards the aft end of the aircraft. These avionics mostly regulate the engine or are “distributed” to improve redundancy, reliability, and durability.

Also, since the avionics components are distributed throughout the aircraft, there will be localized areas that require cooling. These areas should be relatively small, and the heat generated by each group of avionics should be relatively small as well, at least compared to the heat generated by the main laser. Since these small, localized warm spots of avionics groupings are located well within the fuselage, the structures and 38 | P a g e

areas in which they are placed should have the capability of passively cooling the avionics. Essentially, the heat generated by the avionics components is transferred to the interior structure of the fuselage, which absorbs the thermal energy and acts as a heat sink.

Some electronic countermeasures must also be installed to prevent enemies from “hacking” the UAV. This includes features like encrypted signals, redundant systems to ensure the proper functioning of the aircraft, and internal checks on commands.

Figure 7.2.3: The overall avionics placement throughout the fuselage is depicted here. This image contains the avionics described in the Figures 2 and 3 above, with an additional GPS unit and IFF unit towards the middle of the fuselage.

7.3) Communication The major component of the avionics system which allows the UAV to interact with the outside world is the KU SATCOM antenna and data link. This data link allows the UAV to relay digital information to an orbiting military satellite at a rate of 2 MBPS, which is then available to anyone with access to that satellite [16]. This procedure is illustrated in the diagram below, for a Global Hawk UAV in Figure 7.3.1. [17].

This aircraft needs to remain in contact with the satellite, in order to receive orders and transmit information about its environment, status, and mission.

The “launch and recovery element” shown in Figure 7.3.1 above serves as a remote, portable launch site, and must maintain “line of sight” contact with the UAV to continually transmit and

Figure 7.3.1: This diagram depicts the different communications links between a sample UAV (Global Hawk) and support infrastructure and command centers.

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receive information. The long range and long endurance features of this UAV reduce the effectiveness of this possible “launch and recovery element,” since the UAV would likely travel too far before loitering for “line of sight” communications to work well. For this reason, ultra-high frequency (UHF) communications systems were not implemented as part of the UAV’s avionics package.

Likewise, radio communication systems were not necessary, because those systems would mainly be required for communications between the pilot(s) and their base of operations. The data transfer rates for radio signals are too low to be useful (<400 kbps) on a modern UAV, have limited range, and usually require line of sight between the aircraft and receiver to function [18]. Therefore, a radio communication and data transmission system was not included in the avionics package for this UAV.

7.4) Sensors The sensors part of the avionics package includes both sensors required and provided by the RFP and sensors deemed necessary as part of the design process. The sensors required by the RFP are primarily used to detect enemy ICBM missiles in their launch phase, as part of the UAV’s mission objective, and to determine if the targeted ICBM’s have been destroyed after firing the main laser. The other sensors deemed necessary for the successful operation of the UAV include a SAR radar unit, the RWR, a radar altimeter, an IFF receiver unit, and a redundant ILS to aid with landing.

The RFP stipulates that two identical laser sensors must be added to the design of the UAV. These each have a 1 kW power rating, and a 30% electrical-to-optical energy conversion efficiency, and are described in detail in the “Power Requirements” section. One of the laser sensors is used to detect and track the ICBM in the launch phase, as a means of gathering missile position data to aim the 4 MW main laser. The other laser sensor is used to locate disturbances or anomalies in the atmosphere, to assist with both the flight path of the UAV and the targeting of the ICBM missile.

The RFP also stipulates that an infrared sensor must be added to the design of the UAV. This sensor is used to detect launched ICBMs using their heat signatures. This sensor requires 40 kW, and serves as the primary means of identify and tracking targets. This sensor can also be used to help confirm a successful detonation of a targeted missile, since a large burst of infrared radiation followed by a loss of the missile’s heat signature accompanies the destruction of the targeted missile.

Figure 7.4.1: Capabilities of the AN/APG-77 Synthetic Aperture Radar System

The largest and most important of the non-RFP sensors is the radar. A synthetic aperture radar (SAR) type of radar is used in the UAV’s avionics package. The particular radar system selected for this UAV was an AN/APG-77 SAR, also found on the F-22 Raptor. This radar has a range greater than 138 mi (120 nmi) and was selected over similar radar systems due to its longer range, with its specific capabilities depicted in Figure 7.4.1 below [19]).

The range approaches that of the main laser, which has a range of 250 miles. While the infrared sensor and laser sensors should be capable of 40 | P a g e

detecting the target missile at a 250 mi range, this long-range radar system should add redundancy to strikes against missiles within 138 mi of the aircraft. [20] Some sensors are also included as part of the electronic warfare (EW) suite, and the rest are miscellaneous sensors that assist the UAV in some way. The major EW sensor contained within this avionics package is the RWR unit. The RWR has small antennae placed throughout the aircraft, to sense enemy radar “pings” of this UAV. This system is heavily integrated with the flight control computers, to determine the direction of the enemy radar, and with the radar jammer, which can nullify the enemy’s radar capabilities.

Therefore, the RWR system is an important sensor system of the aircraft, although other systems also contribute to the UAV’s situational awareness.

These other miscellaneous sensor systems that assist the UAV include the radar altimeter, the IFF receiver unit, and the instrument landing system (ILS). The radar altimeter uses a downward-directed radar that can determine the altitude of the UAV, which is necessary for flight path considerations, weather conditions, and general aircraft situational awareness. There is also an IFF receiver unit, which senses incoming IFF signals and can distinguish friendly aircraft from enemy aircraft or noise. Finally, the ILS can aid with landing in certain locations, as a backup to either remotely controlled or automated landing procedures. The GPS system would primarily be used to navigate the aircraft during landing operations, and that navigation “sensor” is described in detail elsewhere. Likewise, the KU Satellite antennae “sensor” is also described in more detail in the “Communication” section.

7.5) Flight Control Computers The computer systems selected for the Flight Control and Mission Computers were identical, both being the SRC Mid-Sized Signal Data Processor (SDP) [21]. This particular model of airborne computer was chosen since the maximum allowable altitude for the computers is 60,000 ft, and the UAV will operate between altitudes of 40,000 ft and 60,000 ft.

Other computer systems (including the more powerful SRC ATLAS computer bank) were considered, but either their weight and power requirements were higher than the SRC SDP pair, or they could not operate at the requisite altitudes.

Since one SRC SDP unit will be used to manage the radar alone, a “pair” of two SRC SDP units will serve as a computer “package” or system that satisfies all the processing requirements of the aircraft. The SRC SDP unit not managing the radar will manage the computational processes of the other avionics subsystems. The flight control and mission computers are critical to the UAV’s functioning, so each computing system is triple redundant. The SRC SDP units are Figure 7.5.1: Overview of Systems Architecture for the Control Electronics (ICP, ECC, PMC, PIU) distributed throughout the aircraft to 41 | P a g e

prevent a total loss of computing capability if part of the UAV is damaged or destroyed (see Figure 7.1.1 for Distribution Theory, Figure 7.2.2 for actual location). These SRC SDP computer pairs each have more processing power than necessary for the aircraft’s current avionics loadout and a modular architecture for processing sub-units, so that future extensions or additions to the this avionics package do not require new, more powerful computers [22].

7.6) Power To power the aircraft avionics systems, our design uses the Honeywell APU 131-9[A] with a starter generator that can provide up to 90kW of energy. The power distribution is shown below.

The APU’s 230 VAC powers the aircraft’s larger electrical loads, while multiple rectifier units supply the lower energy devices onboard. The large electrical loads include the infrared sensor module at 40kW, the radar at 12kW, and the fuel pumps. This subset draws a combined power of about 53kW. The transformer rectifier unit (TRU) converts our AC input to DC output. This power passes through a remote power distribution channel to the Flight control computer (FCC), Mission control computer, and the UAV management systems. These processors use about 1000W.

TRU: Transformer Rectifier Unit ATRU: Autotransformer Rectifier Unit RPDU: Remote Power Distr. Unit

Infrared Sensor Module Large Electrical Loads

Radar Fuel Pumps FCC

230 VAC from APU and engine Starter Generators

TRU (28Vdc)

RPDU

MCC

2 1kW Laser Sensors

UAV Management SYstems

Motor Control Devices ATRU (270Vdc) Radar Jammer Sattelite Reciever

Figure 7.6.1: Hierarchy of power distribution.

The autotransformer rectifier unit (ATRU) runs DC output into the two laser sensors, which require 6667W at 30% efficiency as specified by the RFP. The radar jammer can run anywhere from 10 to 4000W, and the satellite receiver requires 825W. These devices use about 12kW. For the flight and mission computers, the TRU excels at electrically isolating the individual processors. This helps prevent shock hazards, noise, and other undesired feedback. However, the ATRU is a better fit for the laser sensors, motor control devices, radar jammer, and satellite receiver due to its higher capacity for power handling. The ATRU also benefits from lower distortion and flatter frequency response.

In the event of an APU failure, the UAS has a battery to provide DC current to the avionics. This battery is sized to allow the UAS to fly home once a critical failure has occurred, from its loiter position. This additional level of redundancy is to minimize the probability of an aircraft loss due to a power system failure.

Upon the UAS switching to battery power, it will automatically return to the base where it is stationed. If the battery begins to fail while returning to its home base, the UAS will be switched over to manual control so that a safe alternate landing site can be found. Because the time it takes to reach the theater of operations is slightly 42 | P a g e

under five hours, the battery was sized to provide an extra hour of flight time with essential avionics, such as the Flight Control Computers and Mission Control Computers, and fuel pumps powered on. The battery weight and volume was calculated using the following equations: đ?&#x2018;&#x192;đ?&#x2018;&#x17D;đ?&#x2018;Łđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018; đ?&#x2018;&#x2021;đ?&#x2018;&#x201C;đ?&#x2018;&#x2122;đ?&#x2018;&#x2013;đ?&#x2018;&#x201D;â&#x201E;&#x17D;đ?&#x2018;Ą = đ?&#x2018;&#x161;đ?&#x2018;?đ?&#x2018;&#x17D;đ?&#x2018;Ąđ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;Ś đ?&#x153;&#x152;đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x201D;đ?&#x2018;Ś đ?&#x2018;&#x192;đ?&#x2018;&#x17D;đ?&#x2018;Łđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018; đ?&#x2018;&#x2021;đ?&#x2018;&#x201C;đ?&#x2018;&#x2122;đ?&#x2018;&#x2013;đ?&#x2018;&#x201D;â&#x201E;&#x17D;đ?&#x2018;Ą = đ?&#x2018;&#x2030;đ?&#x2018;?đ?&#x2018;&#x17D;đ?&#x2018;Ąđ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;Ś đ?&#x153;&#x152;đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x201D;đ?&#x2018;Ś Using the same type of battery as utilized by the laser power system [23], the mass of the battery is 1750 lbs to provide approximately 19.6 kW for 6 hours. The battery has a total volume of approximately 27 ft3.

7.7) System Architecture The system architecture for this project is essentially a hub system, with each avionics component directly linked to the flight and mission control computers. The components that must exchange data for the proper functioning of each subsystem are also connected, just so there are no inconsistencies and proper actions or sequences are performed based upon valid data. The diagram below, Figure 7.5.1, is a graphical depiction of the system architecture for a generic high-altitude long endurance (HALE) UAV, with particular emphasis placed on electronic controllers. This system architecture can be adapted for the UAV described by this report, because specific avionics components comprise each interconnected structure in the system

Figure 7.7.1: Specific Avionics Systems Architecture for the Major Components of the UAV (neglecting the electronic controller architecture depicted in Figure 7 earlier)

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architecture. These components reliably share information with each other using fiber optic cables, which have high data transmission speeds. The above figure accounts for most of the control electronics and shows their connections to each other, their specific controlled devices, and the central flight control and mission control computers. Other systems that are specific to this UAV and its avionics load out still must be addressed from a systems architecting standpoint, and are noted in Figure 7.7.1 below. These systems include Communications, Navigation, and Identification (CNI) features, which are not adequately addressed in Figure 7.5.1 above. These features relate primarily to the sensors and aircraft electronics not encompassed by the electronic controllers.

As depicted by the above figure, Figure 7.7.1, this UAV has a centrally connected mission control computer that monitors sensor inputs to ascertain the completion of mission objectives, and a similarly centrally connected flight control computer which receives sensor input and commands the control surfaces of the aircraft based on this input. The sensors relay information to these computers and each other via fiber optic cable, which is faster and more reliable than other information transmission systems (Fiber Optic Association 9). Some of the sensors and auxiliary electronic systems are wired to exchange information necessary to their functioning, which is also depicted in Figure 7.7.1 above. For example, the infrared sensor and the laser sensor used for tracking the target missile may need to share telemetry data to more accurately fire the main laser. A basic outline of the wired data network with specific avionics components for this UAV is shown in Figure 7.7.1. Laser System.

7.8) Power The laser power supply system consists of a set of supercapacitors and battery packs. After the battery packs fully charge the supercapacitors, the supercapacitors discharge their energy into the laser when a shot needs to be fired. Each battery pack can supply up to one shotâ&#x20AC;&#x2122;s worth of energy and is made up of an arrangement of 3 parallel sets of 4 lithium ion battery modules in series for a total energy of roughly 44MW. The battery is a 24Vdc module from Becket Energy with a current capacity of 42.4Ah, and so the configuration for each pack yields 96Vdc 127.2Ah [19]. We selected this battery because it meets the high average in energy density for lithium ion batteries of today, although other batteries are also appropriate. The design uses 8 battery packs to supply 8 shots for the high-efficiency laser (HEL) and one extra battery pack to power the cooling system. The supercapacitor set consists of 160 supercapacitors from the Maxwell 75V module. The capacitance of each supercapacitor is 94F. This design takes several limitations for energy discharge into consideration. Supercapacitors are ideal because of their ability to quickly discharge energy. A major drawback is that the energy density of supercapacitors is rather low. Even in this design, the energy density is only 2.9Wh/kg which led to a large increase in weight. However, technology around supercapacitors is also still developing and so it is expected for their energy density to increase in the future. The weight required for this design will most likely drop by 2035. Supercapacitors also save costs throughout the 20 year service life of the aircraft. This particular supercapacitor is projected to have a cycle life of around 1,000,000 cycles, so the need for replacement is low.

We used batteries because a generator would be too heavy for our target weight. The overloading capacity of batteries are also more ideal than generators since the overloading the generators can deteriorate the moving parts. It is possible to use lithium air batteries instead lithium ion batteries. Lithium air batteries are currently around 2-3 times the energy density of lithium ion batteries and are projected to have approximately 5-10 times more when the technology has further progression. This allows for more laser blasts at approximately the same or less weight. However, lithium air batteries would need air for its use, and the charge/discharge cycles are currently low. The technology for lithium air must improve before it can be fully implemented. The design for the power system is set up in this sort of hierarchy:

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Figure 7.8.1: Shows the Hierarchy of the Power System where B1B8 denotes the battery packs and 7C and 6C denote the number of super capacitors in that channel.

This hierarchy solves the critical wiring problem if the batteries were to directly charge the capacitors through one direct channel. In order to use only one channel, the amount of current passing through the main wire would be in the kA range which would be far too large, and the wiring resistance alone would have a large impact on the amount of current being drawn. Thus, in this design we parallelize the charging of multiple smaller blocks of capacitors. Each battery pack is intended to represent 8 separate shots of the HEL, however, the design incorporates the use of all of the battery packs at once to implement the parallel charging design.

Each battery pack has 3 parallel channels at 96V, and so 2 of those channels is intended to charge 7 capacitors and 1 is intended to charge 6.

Within each sub section the of each battery pack, the circuit is as follows: The estimates for this circuit were based off the transient RC equation:

Figure 7.8.2: Shows the charging circuit for each block of 7 super capacitors. The ESR for each S.Capacitor is also included.

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đ?&#x2018;Ą

đ?&#x2018;&#x2030;đ??ś = đ?&#x2018;&#x2030;đ?&#x2018;&#x2013;đ?&#x2018;&#x203A; (1 â&#x2C6;&#x2019; đ?&#x2018;&#x2019; â&#x2C6;&#x2019;đ?&#x153;? ) đ?&#x2018;¤â&#x201E;&#x17D;đ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x2019; đ?&#x2018;Ąâ&#x201E;&#x17D;đ?&#x2018;&#x2019; đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;&#x161;đ?&#x2018;&#x2019; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;&#x203A;đ?&#x2018; đ?&#x2018;Ąđ?&#x2018;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018;Ą đ?&#x153;? = đ?&#x2018;&#x2026;đ??ś The voltage and current outputs for this circuit are:

Figure 7.8.4: Shows the current output across each capacitor and the total current across the resistance with respect to time in the 7 capacitor block circuit.

Figure 7.8.3: Shows the Voltage output of each capacitor with respect to time in the 7 capacitor block circuit.

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There were couple requirements that influenced this circuit design. First of all, each supercapacitor is rated at 75V and their low temperature peak continuous current is 48ARMS. This led to laying out the battery modules in a 96V 3 parallel path system, as this allows the 75V to be approximately within 1.53Ď&#x201E;. The value for the resistance here was chosen specifically to meet an approximate time of 10min to charge the supercapacitors. Here the value is 610seconds. This allows a generous time for the cooling systems to cool down the power system after each shot, and keeps the current draw across each capacitor below their 48ARMS limits. The peak value of the current across each capacitor is 22.23A, and the total peak value across the main wire is 155.6A. This allows us to use 2/0 gauge wiring (based off the AWG guide). Also, for the 6 capacitor blocks, the charging time is shorter at approximately 525 seconds and the current spread across each capacitor is 25.81A. The transient voltage graph is shown below for continuity:

Figure 7.8.5: Shows the voltage output of each capacitor with respect to time for the 6 S. capacitor block.

Since the RFP does not specify voltage and load parameters for the HEL, the capacitor discharge process follows standard procedures to disburse 4MW over 10 seconds.

The cooling time for this system can actually be decreased by implementing more battery packs for more shots. However, the only limitation that would need to be kept in mind is the current specifications across each supercapacitor since adding more battery packs would decrease the amount of capacitors in each channel.

This increases the current passing through. The 8 shots for this system and the 10 minute timing prioritize the cooling of the laser system.

The laser power system incorporates one extra battery pack for the cooling system of the laser. The estimated energy usage of the cooling system is 1.8kW. Its worst case, at around 10min of operation for 8 shots of the laser, is an energy use of approximately 8.64MW. Thus, the 44MW pack will be sufficient.

The battery system alone also incorporates a deployment structure that makes it easy to detach from the aircraft. Its total weight is supportable with current aircraft standards. This will allow the batteries to be easily recharged off the aircraft, and allows for easy replacement if necessary.

7.9) Cooling 7.9.1) Background Convective heat transfer is governed by newtonâ&#x20AC;&#x2122;s law of cooling. đ?&#x2018;&#x2018;đ?&#x2018;&#x201E; = â&#x201E;&#x17D; â&#x2C6;&#x2014; đ??´(đ?&#x2018;&#x2021; â&#x2C6;&#x2019; đ?&#x2018;&#x2021;đ?&#x2018;?đ?&#x2018;˘đ?&#x2018;&#x2122;đ?&#x2018;&#x2DC; ) đ?&#x2018;&#x2018;đ?&#x2018;Ą 47 | P a g e

â&#x201E;&#x17D; = â&#x201E;&#x17D;đ?&#x2018;&#x2019;đ?&#x2018;&#x17D;đ?&#x2018;Ą đ?&#x2018;Ąđ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018; đ?&#x2018;&#x201C;đ?&#x2018;&#x2019;đ?&#x2018;&#x; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;&#x2019;đ?&#x2018;&#x201C;đ?&#x2018;&#x201C;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;Ą đ??´ = đ?&#x2018;¤đ?&#x2018;&#x2019;đ?&#x2018;Ąđ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x2018; đ?&#x2018;&#x17D;đ?&#x2018;&#x;đ?&#x2018;&#x2019;đ?&#x2018;&#x17D; đ?&#x2018;&#x2021; = đ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x161;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;Ąđ?&#x2018;˘đ?&#x2018;&#x;đ?&#x2018;&#x2019; It is difficult to solve this equation but estimations can be made to the convective heat transfer coefficient if we know certain material and dynamic properties. The convective heat transfer coefficient, h, is related to the Nusselt Number, the ratio of convective to conductive heat transfer normal to a boundary. Nu and can be calculated as follows. đ?&#x2018;&#x2DC; â&#x201E;&#x17D; = đ?&#x2018; đ?&#x2018;˘ đ??ˇ đ?&#x2018;&#x2DC; = đ?&#x2018;Ąâ&#x201E;&#x17D;đ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x161;đ?&#x2018;&#x17D;đ?&#x2018;&#x2122; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;&#x203A;đ?&#x2018;&#x2018;đ?&#x2018;˘đ?&#x2018;?đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;Łđ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś đ??ˇ = â&#x201E;&#x17D;đ?&#x2018;Śđ?&#x2018;&#x2018;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;˘đ?&#x2018;&#x2122;đ?&#x2018;&#x2013;đ?&#x2018;? đ?&#x2018;&#x2018;đ?&#x2018;&#x2013;đ?&#x2018;&#x17D;đ?&#x2018;&#x161;đ?&#x2018;&#x2019;đ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x; The Nusslet number depends on many factors but it can be approximated under turbulent flow using the Gnielinski correlation. đ?&#x2018;&#x201C; ( ) (đ?&#x2018;&#x2026;đ?&#x2018;&#x2019; â&#x2C6;&#x2019; 1000)đ?&#x2018;&#x192;đ?&#x2018;&#x; 8 đ?&#x2018; đ?&#x2018;˘ = 1 2 đ?&#x2018;&#x201C; 2 1 + 12.7 ( ) ((Pr)3 â&#x2C6;&#x2019; 1) 8 đ?&#x2018;&#x2026;đ??¸ = đ?&#x2018;&#x2026;đ?&#x2018;&#x2019;đ?&#x2018;Śđ?&#x2018;&#x203A;đ?&#x2018;&#x153;đ?&#x2018;&#x2122;đ?&#x2018;&#x2018;đ?&#x2018; đ?&#x2018; đ?&#x2018;˘đ?&#x2018;&#x161;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x; Pr = đ?&#x2018;&#x192;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018;&#x2018;đ?&#x2018;&#x2122;đ?&#x2018;Ą đ?&#x2018; đ?&#x2018;˘đ?&#x2018;&#x161;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x; đ?&#x2018;&#x201C; = đ??ˇđ?&#x2018;&#x17D;đ?&#x2018;&#x;đ?&#x2018;?đ?&#x2018;Ś đ?&#x2018;&#x201C;đ?&#x2018;&#x;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A; đ?&#x2018;&#x201C;đ?&#x2018;&#x17D;đ?&#x2018;?đ?&#x2018;Ąđ?&#x2018;&#x153;đ?&#x2018;&#x; đ?&#x2018;&#x2030;đ?&#x2018;&#x17D;đ?&#x2018;&#x2122;đ?&#x2018;&#x2013;đ?&#x2018;&#x2018; đ?&#x2018;&#x2013;đ?&#x2018;&#x201C;: 0.5 â&#x2030;¤ đ?&#x2018;&#x192;đ?&#x2018;&#x; â&#x2030;¤ 2000, 3000 â&#x2030;¤ đ?&#x2018;&#x2026;đ??¸ < 5 â&#x2C6;&#x2014; 106 The Reynolds Number can be calculated as follows: đ?&#x153;&#x152;đ?&#x2018;&#x2030;đ??ˇ đ?&#x2018;&#x2026;đ??¸ = đ?&#x153;&#x2021; đ?&#x153;&#x152; = đ?&#x2018;&#x2018;đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018; đ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś đ?&#x2018;&#x2030; = đ?&#x2018;&#x161;đ?&#x2018;&#x2019;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A; đ?&#x2018;Łđ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;Ąđ?&#x2018;Ś đ?&#x2018;&#x153;đ?&#x2018;&#x201C; đ?&#x2018;&#x201C;đ?&#x2018;&#x2122;đ?&#x2018;˘đ?&#x2018;&#x2013;đ?&#x2018;&#x2018; đ?&#x153;&#x2021; = đ??ˇđ?&#x2018;Śđ?&#x2018;&#x203A;đ?&#x2018;&#x17D;đ?&#x2018;&#x161;đ?&#x2018;&#x2013;đ?&#x2018;? đ?&#x2018;&#x2030;đ?&#x2018;&#x2013;đ?&#x2018; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018; đ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś The Prandlt number can be calculated as follows: đ?&#x2018;&#x192;đ?&#x2018;&#x; =

đ??śđ?&#x2018;? đ?&#x153;&#x2021; đ?&#x2018;&#x2DC;

The Darcy Friction Factor can be estimated using the BrkiÄ&#x2021; solution: 1 đ?&#x153;&#x20AC; 2.18 â&#x2C6;&#x2014; đ?&#x2018;&#x2020; = â&#x2C6;&#x2019;2 log10 ( + ) 3.71 â&#x2C6;&#x2014; đ??ˇ đ?&#x2018;&#x2026;đ??¸ â&#x2C6;&#x161;đ?&#x2018;&#x201C; đ?&#x2018;&#x2026;đ??¸ ) 1.1 â&#x2C6;&#x2014; đ?&#x2018;&#x2026;đ??¸ 1.816 â&#x2C6;&#x2014; đ?&#x2018;&#x2122;đ?&#x2018;&#x203A; ( ) đ?&#x2018;&#x2122;đ?&#x2018;&#x203A;(1 + 1.1 â&#x2C6;&#x2014; đ?&#x2018;&#x2026;đ??¸ đ?&#x153;&#x20AC; = đ?&#x2018;&#x;đ?&#x2018;&#x153;đ?&#x2018;˘đ?&#x2018;&#x201D;â&#x201E;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018;&#x2019;đ?&#x2018; đ?&#x2018; * đ?&#x153;&#x20AC; đ?&#x2018;&#x2013;đ?&#x2018; đ?&#x2018;&#x17D;đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;˘đ?&#x2018;Ą .002 đ?&#x2018;&#x161;đ?&#x2018;&#x161; đ?&#x2018;&#x201C;đ?&#x2018;&#x153;đ?&#x2018;&#x; đ?&#x2018;&#x17D;đ?&#x2018;&#x2122;đ?&#x2018;˘đ?&#x2018;&#x161;đ?&#x2018;&#x2013;đ?&#x2018;&#x203A;đ?&#x2018;˘đ?&#x2018;&#x161;

đ?&#x2018;&#x2020; = ln (

Given these relations, the goal would be to maximize the surface area while still retaining turbulent flow.

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7.9.2) Laser Housing Sizing By maximizing the surface area, this will enable the laser system to radiate more heat off of its surfaces. Given the constraints of the laser sizing, a MATLAB script was used to find the maximum surface area. The MATLAB script iterates through different sizes for the width (X) and the height (Y) while keeping the volume (V) constant at 706.3 m3, then calculating the surface area. This max surface area is then reported along with the parameters for the length, height, and width. Based on this analysis, weâ&#x20AC;&#x2122;ve sized out the following dimensions for the laser Table 7.9.1: Laser Dimensions

Length 10.5 Feet

Width 8.2 Feet

Height 8.2 Feet

7.9.3) Channel Sizing To cool off the laser, coolant will be running through square cooling channels. One side of the cooling channels will be exposed to the heat. With square cooling channels, the system would maintain better contact with the heated area and would be able to dissipate heat more efficiently.

At altitudes of 40,000 ft., the average air temperature is approximately -70 °F, or -57 °C. Taking this into account with the lack of cabin control in the fuselage, weâ&#x20AC;&#x2122;ve estimated that the initial temperature of the fluid and the cooling system would be about -40 degrees F, or -40 degrees C. This means that for JP-8, the density would be 850 kg/m3 or 53 lb./ft.3 The 8000 dynamic viscosity at this temperate would be approximately 0.12 Pa-s or 1.2 poise.

7000 The dynamic viscosity is rather high, so a 6000 velocity of 1 meter per second, or 3.28084 feet per second, has been chosen in order 5000 to maintain turbulence despite the high 4000 dynamic viscosity.

For the sizing of the channels, 3000 first some initial CFD results with different geometries and flow rates, it was 2000 determined that the height of the channels 1000 should be 1.25 inches while being .2 inches thick. This allows the heat to travel 0 up the walls while minimizing weight.

0 20 40 60 80 The Reynolds number for the flow should # of channels be a little bit above 4000 so weâ&#x20AC;&#x2122;re still in the turbulent region while exposing as Figure 7.9.1: Number of Channels vs. RE much surface area as we can. The width of the channel can be calculated using the following formula. (đ?&#x2018;Ąđ?&#x2018;¤đ?&#x2018;&#x17D;đ?&#x2018;&#x2122;đ?&#x2018;&#x2122; â&#x2C6;&#x2014; (# đ?&#x2018;&#x153;đ?&#x2018;&#x201C; đ?&#x2018;?â&#x201E;&#x17D;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018;&#x203A;đ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018; + 1)) đ?&#x2018;&#x160;đ?&#x2018;?â&#x201E;&#x17D;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A; = (đ?&#x2018;&#x160;đ?&#x2018;&#x153;đ?&#x2018;Łđ?&#x2018;&#x2019;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;&#x2122;đ?&#x2018;&#x2122; â&#x2C6;&#x2019; # đ?&#x2018;&#x153;đ?&#x2018;&#x201C; đ?&#x2018;?â&#x201E;&#x17D;đ?&#x2018;&#x17D;đ?&#x2018;&#x203A;đ?&#x2018;&#x203A;đ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018; Combining this with the formula for Reynolds number, a relationship has been defined. Figure 8.2.1: Number of Channels vs. RE shows how the Reynolds number changes with # of channels. Ideally the RE

# of Channels vs RE

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Reynolds number should be around 4000 because this is the transition point between fully laminar and fully turbulent flow. Additionally, the channels should have a somewhat round number to help with manufacturing tolerances. Based on the calculations, the number of channels that have been selected is 50, making the width of each channel 1.764 inches, and a Reynolds number of 4400.

7.9.4) Fluid Dynamics Simulations To get an idea of the performance of the cooling channels, Solidworks Flow Simulation was used to run studies using computational fluid dynamics. Solving for the entire system is too computationally intensive, so the simulation has been scaled down to a group of 5 channels together, keeping the same channel geometries and surface heat flux. The surface heat flux was calculated by taking the total heat dissipation (3000000 W) and dividing it by the total amount of surface area being exposed (8.2 feet * 10.5 feet * 4 = 344.4 feet or 32 meters). This means the total heat flux would be 93762.30 W/m2. Figure 8.2.2: Temperature vs time below shows the laser temperature as a function of time for the basic straight channel with the given channel geometry (1.764 inch by 1.25 inch). The timescale for this will be 10 seconds, simulating a 10 second firing time for the laser. The temperature reported corresponds to laser temperature.

The graph above estimates the laser temperature by calculating the bulk average temperature of the heated surface. With this cooling system in place, the laser reaches an average temperature of 4.74 째C, with a max local temperature of 7.646 째C.

Temperature vs Time Temperature (Solid) [째C]

10 0 -10

0

2

4

6

8

10

12

-20 -30 -40

-50

Physical time (s) Figure 7.9.2: Temperature vs time

7.9.5) Passive mixing using ramps

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Heat will be exposed to one side of the cooling channels, which would mean there would be a clear thermal gradient along the z direction of the flow. This could cause issues if local pockets of fluid reached their boiling points, especially at the bottom corners where heat transfer is the greatest. See Figure 8.2.3: Temperature Gradient below for the thermal gradient.

Figure 7.9.3: Temperature Gradient

In order to promote better mixing and possibly convective heat transfer, circular ramps have been added to the channel. The ramps have been angled to allow the fluid to swirl around the channel. This form of passive mixing should make the temperature distribution much more even across the field. In addition, the ramps will help create a turbulent wake across the channel, improving convective heat transfer. See the figures below for how they are configured.

Figure 7.9.4: Streamlines

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Figure 7.9.6: top view of the ramps

Figure 7.9.5: side view of the ramps

The ramps are spaced in such a way so the turbulent wake left behind isnâ&#x20AC;&#x2122;t disturbed, allowing for greater heat transfer in that localized region. The turbulent wake has been determined to be about 5 inches according to the CFD results. The ramps are also slanted in so the fluid flow would be redirect vertically and horizontally so this encourages mixing as the fluid flows down. Figure 8.2.4: Streamlines shows the path of the fluid as a result of the ramps.

The presence of these ramps causes a noticeable difference in how the thermal gradient at the exit of the channel. Figure 8.2.7 below shows a much smoother temperature contour as a result.

When we run CFD though a group of 5 of these channels together, improvements are small but noticeable.

The average temperautre of the laser after 10 seconds of continuous firing was 4.06 degrees C. The max local temperatre observed was 8.01 degrees C. This shows abut a 1 degree improvement over straight channels.

Figure 7.9.7: Temperature Gradient with ramps

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7.9.6) Routing through multiple channels Since heat transfer should improve if the coolant has been heated up, it might be advantageous to have the fluid route through multiple fluid channels before exiting. This way, the coolant has more time to heat up, thus improving heat transfer as it travels though the cooling channel. For this simulation, the fluids will route through 5 channels before exiting. However, with this configuration, large pressure losses are expected. See figure 8.2.8 for the model.

With this in mind, the simulation was run using the same settings as the previous simulation. Fluid velocity was set to 1 m/s, or 3.28 ft/s. Heat flux for the heated surface stayed the same, 93762.30 W/m2.

A clear temperature difference is noticeable as the fluid travels from the right channel to the left channel. The contours show decent levels of mixing as the fluid travels down. Figure 8.2.9 below shows the thermal gradients across all 5 channels.

The laser temperature at the end of the firing time was only 3 degrees C, with a local maximum temperature of 7. These results show a lot of promise, as theoretically we should be able to fire off the laser for longer than 10 seconds.

With this in mind, the laser temperature performance should have improved. Below are the results of the thermal simulation

7.9.7) Comparison of channel geometries Simple straight cooling channels provide good cooling performance of the laser system. The max temperature of the laser never goes over 8 degrees C. There is an improvement when the mixing ramps are added to the assembly. Making the fluid loop through the channels provides some performance improvement at the cost of a much larger pressure drop. Despite the large pressure drop and seemly small performance gain, the cooling channels will use the multiple channel design. Temperature of the laser system takes precedence since beam quality depends on the temperature being as low as possible.

Figure 7.9.8: the model for the cooling channels. The fluid entrances and exits are highlighted in blue

53 | P a g e

Therefore, the third design will be used for the entire system.

7.9.8) Placement and Configuration As the laser system itself is a black box, we can only assume that heat will be dissipated through the surface of the box. Therefore, the full size cooling channels will be placed on the surface of the laser box.

Manifolds will be attached to the front and back of each cooling channel in order to split the flow into the channels, and merge them back together. One stream will be flowing through the top cooling, and have it flow back into the adjacent side cooling channel. The other fluid stream will enter the bottom of the side cooling channel, and then loop around into the bottom cooling channel. See Figure 8.2.9: Isometric view for the overall system. Note that fluid will enter through the highlighted parts in the upper corners of the system.

From the front, the piping makes the fluid loop around into the adjacent lower manifold. The piping makes a 90 degree turn to the side before making another 90 degree turn to the next manifold. This is done to leave the front of the laser exposed. The laser will be firing in the front, and this will ensure nothing will be obstructing the beam.

Figure 7.9.9: Temperature contours at the middle of the channel.

7.9.9) Power Requirements Power requirements for the system will be based the amount of power needed to loop the fuel from the tanks, through the cooling system and back into the tanks. This will be a closed loop system, and thus there will be no velocity, pressure, or static head losses. The only energy losses that will occur are frictional losses.

Major losses are attributed to frictional forces within the fluid. This can be calculated as follows. đ?&#x2018;&#x2122; 1 â&#x201E;&#x17D;đ?&#x2018;&#x161;đ?&#x2018;&#x17D;đ?&#x2018;&#x2014;đ?&#x2018;&#x153;đ?&#x2018;&#x; = đ?&#x2018;&#x2DC; ( ) ( ) đ?&#x2018;&#x2018;â&#x201E;&#x17D; 2đ?&#x2018;&#x201D; đ?&#x2018;&#x2DC; = đ?&#x2018;&#x201C;đ?&#x2018;&#x;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;&#x2019;đ?&#x2018;&#x201C;đ?&#x2018;&#x201C;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;Ą đ?&#x2018;&#x2018;â&#x201E;&#x17D; = â&#x201E;&#x17D;đ?&#x2018;Śđ?&#x2018;&#x2018;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;˘đ?&#x2018;&#x2122;đ?&#x2018;&#x2013;đ?&#x2018;? đ?&#x2018;&#x2018;đ?&#x2018;&#x2013;đ?&#x2018;&#x17D;đ?&#x2018;&#x161;đ?&#x2018;&#x2019;đ?&#x2018;Ąđ?&#x2018;&#x2019;đ?&#x2018;&#x; đ?&#x2018;&#x201D; = đ?&#x2018;&#x17D;đ?&#x2018;?đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018;&#x2019;đ?&#x2018;?đ?&#x2018;&#x;đ?&#x2018;Ąđ?&#x2018;&#x17D;đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A; đ?&#x2018;&#x2018;đ?&#x2018;˘đ?&#x2018;&#x2019; đ?&#x2018;Ąđ?&#x2018;&#x153; đ?&#x2018;&#x201D;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;Łđ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś đ?&#x2018;&#x2122; = đ?&#x2018;?đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;&#x2019; đ?&#x2018;&#x2122;đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;&#x201D;đ?&#x2018;Ąâ&#x201E;&#x17D; Not including the manifolds and the cooling channels, the piping length is approximately 220 inches, or 5.6 meters.

The hydraulic diameter is just the diameter of the pipe pipes, which is approximately 3.7 inches, or 9.4 centimeters. The area of this corresponds to 5 times the cross sectional area of one Figure 7.9.9: Isometric view

54 | P a g e

channel, as each manifold splits the fluid into 5 inlets, as this will keep the mass flow rate constant. The friction coefficient, k, depends on the Reynolds number and the pipe roughness. With a hydraulic diameter of 3.7 inches, RE is approximately 68000. The pipe is made of aluminum alloy, so a surface roughness of .002 mm can be assumed. Relative roughness is nearly zero so this mimics the behavior of a smooth pipe. Factoring all of this together, this corresponds to a friction factor of about 0.017. The total major head loss is equal to 1.8 feet or 0.54 meters.

Minor losses are those that are incurred within system components. Minor losses can be calculated as such đ?&#x2018;Ł2 â&#x201E;&#x17D;đ?&#x2018;&#x161;đ?&#x2018;&#x2013;đ?&#x2018;&#x203A;đ?&#x2018;&#x153;đ?&#x2018;&#x; = đ?&#x2018;&#x2DC; ( ) 2đ?&#x2018;&#x201D; đ?&#x2018;&#x2DC; = đ?&#x2018;&#x161;đ?&#x2018;&#x2013;đ?&#x2018;&#x203A;đ?&#x2018;&#x153;đ?&#x2018;&#x; đ?&#x2018;&#x2122;đ?&#x2018;&#x153;đ?&#x2018; đ?&#x2018; đ?&#x2018;?đ?&#x2018;&#x153;đ?&#x2018;&#x2019;đ?&#x2018;&#x201C;đ?&#x2018;&#x201C;đ?&#x2018;&#x2013;đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x203A;đ?&#x2018;Ą đ?&#x2018;Ł = đ?&#x2018;&#x201C;đ?&#x2018;&#x2122;đ?&#x2018;˘đ?&#x2018;&#x2013;đ?&#x2018;&#x2018; đ?&#x2018;Łđ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018;&#x153;đ?&#x2018;?đ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś đ?&#x2018;&#x201D; = đ?&#x2018;&#x17D;đ?&#x2018;?đ?&#x2018;?đ?&#x2018;&#x2019;đ?&#x2018;&#x2122;đ?&#x2018;&#x2019;đ?&#x2018;?đ?&#x2018;&#x;đ?&#x2018;Ąđ?&#x2018;&#x17D;đ?&#x2018;Ąđ?&#x2018;&#x2013;đ?&#x2018;&#x153;đ?&#x2018;&#x203A; đ?&#x2018;&#x2018;đ?&#x2018;˘đ?&#x2018;&#x2019; đ?&#x2018;Ąđ?&#x2018;&#x153; đ?&#x2018;&#x201D;đ?&#x2018;&#x;đ?&#x2018;&#x17D;đ?&#x2018;Łđ?&#x2018;&#x2013;đ?&#x2018;Ąđ?&#x2018;Ś Not counting the manifolds or the cooling channel itself, minor losses 10 are incurred within 90 degree bends in the piping. The minor loss coefficient 0 for a 90 degree elbow is 0.3 [24]. Given 0 2 4 6 8 10 12 the velocity, the minor loss coefficient, hminor, is about .4 feet or .122 meters.

-10 For losses in the manifold and the cooling channel, fluid simulations -20 have been used to observe the pressure loss across the system component. The -30 average pressure loss was calculated to be 0.11 psi or 780 Pascals. Dividing pressure loss by the density and -40 acceleration due to gravity, a head loss of the manifold, the pressure head has -50 been calculated as .30 ft. or 0.09 m.

Physical time (s) Multiplied by 4 per cooling loop, the Figure 7.9.10: Temperature vs Time graph. total head loss is 1.19 feet or .36 m For flow across the channels, pressure loss will be great due to the complex geometry. According to the CFD analysis, the pressure drop was observed to be quite large, 18000 Pascals or 2.6 psi. This corresponds to a head loss of 2.16 m / 7.08 feet per channel, or 4.32 / 14.17 per loop.

With this in mind, we can calculate the total power required. If a 70% efficiency rating is assumed, the total power required is about 2.36 HP, or 1764 W. Temperature (Solid) [°C]

Laser Temperature vs Time

Table 7.9.2: Summary of Cooling System Efficiency

Losses Minor Loss Head Major Loss Head Manifold head Cooling Channel Head Total System Head

English 1.800 ft.

0.549 ft.

1.193 ft.

13.68 ft.

17.08 ft.

SI .122 m .548 m 0.364 m 4.17 m 5.20 m 55 | P a g e

Mass Flow Rate Efficiency Total Power Required

53.31 lb./s 0.7 % 2.36 hp.

24.184 kg/s 0.7 % 1764 W

7.10) Layout The laser diode is placed 58.33 ft from the turret mounting plate and is centered in the circular cross section of the fuselage. The laser is held in place by structures that support the laser cooling system. Because the cooling system is directly attached to the diode, the cooling system supports the diode structurally. The battery pack and capacitor pack are placed at 69.166 ft from the turret mounting plate. The capacitor pack is 2 ft above the battery pack to provide room for wiring and cooling through dissipation of heat to the ambient air surrounding the components. The battery pack is placed approximately flush with the bottom of the cooling system to allow ground crews to easily swap out battery packs to reduce sortie turnaround time. The capacitor pack is permanently attached to the airframe due to its size, weight, and long cyclic lifespan.

8) Materials Selection 8.1) Structure: 2024 – T3 (Flat sheet) Aluminum For the structure of the aircraft, 2024 – T3 aluminum has been selected. This alloy is typically used in aerospace structures such as the fuselage, wing tension members, and sheer webs. 2024 Aluminum exhibits great stiffness, fatigue performance and good strength. In addition, the T3 temper of Aluminum is noted to have excellent fracture toughness as well [25].

The 1xxx series of aluminum alloys are essentially pure aluminum, and would be too soft with insufficient strength for structural applications in aerospace. For other common aerospace alloys, such as 6061, 6063, and 5052, the strength of these materials does not compare to the strength of 2024 – T3 or 7075 – T3.

Table 8.1.1: Strength data obtained from McMaster-Carr for unpolished sheets

Grade / Temper 7075 - T651 6061 T6511

Yield Strength (ksi) 61-68 35

2024 T-3 5052 H38

45 28

When comparing to between 2024 T3 and 7075 – T6, 7075 has a much higher yield strength compared to 2024. However, 7075 – T651 is more sensitive to notches and has a higher fatigue-crack propagation rate. 2024 –T3 will be chosen for its better resistance to wear and tear. [26]

8.2) Cooling channels: 6063-T5 Aluminum Although copper has better thermal properties, it’s too expensive and dense to use (about 3-4 times denser than aluminum). Aluminum will be considered for the cooling channels because it’s very light and has good strength properties. Pure aluminum would be the idea choice for the cooling system piping and heat spreaders. However, it’s quite soft and its strength is rather lackluster. Instead, the cooling channel/heat spreader and pipes will be using 6061 Aluminum Alloy for its better than average thermal characteristics 56 | P a g e

compared to other alloys, excellent corrosion resistance to atmospheric and sea conditions, as well as its good weldability and workability. In addition, this is the least expensive of the Aluminum alloys.

6063 alloy isnâ&#x20AC;&#x2122;t quite as strong as other aluminum alloys but it still exhibits good strength and weight characteristics. We can expect that the cooling channels will not be experiencing the same magnitude of loads that the aircraft structure will face. Thermal properties are more important, and 6063-O has a higher specific heat (0.214961 Btu/F) and thermal conductivity (0.00279 Btu/in-sec-F)

8.3) Cooling Liquid: JP-8 In addition to being a Jet fuel, JP-8 has also been used as a coolant for jet engines and other components [27]. In the same way, JP-8, will be used as the cooling fluid for the laser. This allows us to reduce complexity and weight by not having to account for additional tanks and liquids.

It has been noted that the thermal conductivity of JP-8 decreases as the temperature increases. As the temperature of the fluid increases, thermal conductivity, specific heat, and density decrease. However viscosity decreases as well. Factoring these variables into the system, we see that as fluid temperature increases, an overall increase in the Reynolds number occurs, indicating an increase in turbulence.

Temperature vs RE 14000 12000 10000 8000 6000 4000 2000

0 200

250

300

350

400

Figure 8.3.1: Reynolds Number as a Function of Temperature Table 8.3.1: JP-8 Properties vs Temperature

T (K)

đ?&#x153;&#x152; (kg/m^3)

đ?&#x153;&#x2021; ( Pa/s)

D (m)

V (m/s)

RE

233 243

850 840.7692385

0.011764706 0.007136322

0.063813933 0.063813933

0.2 0.2

922.1113 1503.654

253 263

834.6153923 828.4615462

0.005391705 0.004224698

0.063813933 0.063813933

0.2 0.2

1975.631 2502.777

273 283

822.3077 816.1538538

0.003161833 0.002450518

0.063813933 0.063813933

0.2 0.2

3319.257 4250.691

293 303

810.0000077 803.8461615

0.002160494 0.001803828

0.063813933 0.063813933

0.2 0.2

4784.951 5687.526 57 | P a g e

313

797.6923154

0.001504339

0.063813933

0.2

6767.606

323 333

791.5384692 785.3846231

0.001263362 0.001145935

0.063813933 0.063813933

0.2 0.2

7996.309 8747.174

343 353

779.2307769 773.0769308

0.001026653 0.000944279

0.063813933 0.063813933

0.2 0.2

9686.965 10448.84

363 770 0.000831169 0.063813933 0.2 11823.53 Noting this, we also need to consider the Prandlt number (Pr), the ratio of momentum diffusivity (kinematic viscosity) to thermal diffusivity. The heat transfer coefficient is also a function of the Prandlt number since the coefficient of heat transfer and the thermal conductivity depend on the temperature. As shown below, as the temperature of the fluid increases, Pr decreases.

Table 8.3.2: JP-8 Prandtl Number as a Function of Temperature

T Cp u k Pr 233 1625 0.011765 0.1264 151.2472 243 1675 0.007136 0.1244 96.08794 253 1725 0.005392 0.1224 75.98604 263 1775 0.004225 0.1204 62.28272 273 1825 0.003162 0.11875 48.59239 283 1875 0.002451 0.1165 39.43967 293 1925 0.00216 0.11505 36.14907 303 1975 0.001804 0.113125 31.49224 313 2025 0.001504 0.11175 27.25984 323 2075 0.001263 0.11 23.83161 333 2125 0.001146 0.10875 22.39184 343 2175 0.001027 0.10675 20.91776 353 2225 0.000944 0.1024 20.51777 363 2275 0.000831 0.1004 18.83376 Although Pr decreases with increasing T, the increased presence of turbulent flow is enough to overcome the decrease in thermal conductivity, meaning the heat transfer coefficient actually increases as T increases. The friction factor f, and hydraulic diameter D are fixed. Thus, we can expect consistent or better

Tvh 290

h (w/m^2k)

240 190 140 90 40 -10

200

250

300 T (k)

350

Figure 8.3.2: Conductive Heat Transfer Coefficient vs. Temperature

400 58 | P a g e

cooling performance as the fluid is being heated.

Table 8.3.3: JP-8 Cooling Performance as a Function of Temperature

T 233 243 253 263 273 283 293 303 313 323 333 343 353 363

Re 922.1113 1503.654 1975.631 2502.777 3319.257 4250.691 4784.951 5687.526 6767.606 7996.309 8747.174 9686.965 10448.84 11823.53

Pr 151.2472 96.08794 75.98604 62.28272 48.59239 39.43967 36.14907 31.49224 27.25984 23.83161 22.39184 20.91776 20.51777 18.83376

k 0.1264 0.1244 0.1224 0.1204 0.11875 0.1165 0.11505 0.113125 0.11175 0.11 0.10875 0.10675 0.1024 0.1004

f 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776 0.053776

D 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814 0.063814

Nu -2.68279 14.92061 26.73625 38.55291 54.79941 71.67778 81.08604 95.94384 112.5513 130.6009 141.6732 155.33 167.8805 186.9513

h -5.31396 29.08649 51.28217 72.73913 101.9751 130.8564 146.1899 170.0827 197.0982 225.1248 241.4357 259.8411 269.3921 294.135

Pr

Pr vs T 160 140 120 100 80 60 40 20 0 200

250

300

350

400

Temperature (K) Figure 8.3.3: Prandlt Number as a Function of Temperature

8.4) Nozzle: Grade 6 Ti Alloy Titanium alloy has been selected for the nozzle material. Titanium, like aluminum, has a high strength to density ratio and great corrosion resistance. In fact the strength for titanium alloys are much greater than the strength of aluminum alloys. Additionally, titanium alloys exhibit excellent fatigue strength and life in the air, unaffected by most other corrosive environments like seawater [28]. Also, aluminum alloys show great resistance to creep at elevated temperatures, important for jet nozzles.

Grade 6 titanium is a commonly used titanium alloy in the aerospace industry, particularly for gas turbine components. This alloy is most commonly used for applications requiring good weldability, stability and strength at elevated temperatures [29]. Additionally, it shows good oxidation and creep resistance.

59 | P a g e

8.5) Wings and skin: Woven AS4 Carbon Fiber Being much lighter than metals, Carbon fiber would be the ideal material for composites. Carbon fiber has a density of only 0.0647 lb/in3 while maintain a tensile yield strength of 320 ksi when it is used to form a composite [30]. Compare that to typical alloys such as aluminum 6061 (density of 0.0975 lb/in3, tensile yield strength of 40 Ksi), or grade 5 titanium (density of 0.16 ln/in3, tensile yield strength of 130 ksi), carbon fiber is the ideal material to reinforce the aircraft structure.

Woven fiber sheets will be selected instead of unidirectional ones as woven fibers are better at handling loads that are off axis. This especially helps in less than idea situations, such as when the fibers have been punctured or it is exposed to bearing loads. [31]

8.6) Leading Edge of Wings: 2124 – T851 The leading edge of the wings will be made out of 2124 – T851 aluminum alloy. Aside from having a high strength to weight ratio and great fatigue performance, the T851 temper of aluminum has great resistance to Stress Corrosion Cracking (SCC) and creep resistance at higher temperatures (up to 350 degrees F). In over 20 years of service experience there have been no reported incidents of SCC failures in 2024-T351 or 2024T851 materials [25]. In addition, military specifications for 2124-T851 alloy state that there should not be any stress corrosion problems with this alloy. Test specimens passed stress corrosion tests for the short transvers e (S-T) orientation. [32]

8.7) Radar Dome: Fiber Glass Standard fiber glass will be used for the laser radar dome. Fiber glass is lightweight, has a high strength, and is easy to form. The tensile yield strength of fiberglass and polyester is about 30 ksi, which is comparable to 6061 T-6 aluminum (35 ksi). However it’s noted that the material cost per lb for fiberglass is only $1.80 / lb which is almost ½ that of aluminum 6061, $3/lb [33]. In addition, its density is almost ½ that of Aluminum 6061-T6 (0.055 in/lb3 vs .10 lb/in3). This makes fiber glass ideal in terms of mechanical properties.

In addition, fiber glass also has a high RF permeability, low signal attenuation properties, which is critical in communications systems. This material has been a standard material for making radomes since World War 2.

8.8) Laser Housing: Duroplast Thermoset The laser housing will be composed of duroplast, which has high stiffness and hardness, low tendency to creep, high heat forming resistance, low thermal linear expansion, and low flammability. This makes it idea as a housing material for the laser as we can expect high thermal dissipation within the housing. In addition, duroplast can also be used to thermally isolate sensitive components [34].

9) Dynamics and Controls 9.1) Un-augmented Dynamics An aircraft’s dynamics are typically separated into two parts for reduced complexity when calculating the aircrafts response: the longitudinal axis and the lateral-directional axis. The longitudinal axis consists of the aircraft’s response to elevator deflections and changes in engine throttle. The lateral-directional axis contains the aircraft’s response to aileron and rudder deflections. To model the dynamics of the UAS, linearized versions of the equations of motion were utilized to facilitate analysis of the dynamics. The resulting linearized state space equations are:

60 | P a g e

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đ?&#x2018;&#x152;đ?&#x203A;żđ?&#x2018;&#x; đ?&#x2018;˘0 Î&#x201D;đ?&#x203A;ż đ??żđ?&#x203A;żđ?&#x2018;&#x; [ đ?&#x2018;&#x17D; ] Î&#x201D;đ?&#x203A;żđ?&#x2018;&#x; đ?&#x2018; đ?&#x203A;żđ?&#x2018;&#x; 0 ]

These equations were input into MATLAB utilizing the built in state space variable type with a four by four identity matrix for C and an empty matrix for D so the dynamics of the aircraft could be analyzed using Bode or Nyquist plots and eigenvalue analysis. The values for the stability derivatives and parameters were determined using Advanced Aircraft Analysis (AAA). Because AAA uses mathematical models for determining the stability derivatives and parameters, the numbers are an estimate for the actual dynamics of the aircraft which would have to be determined through vigorous wind tunnel and ground testing of scale models of the proposed aircraft design. However as conducting such a test program is well beyond the resources of this design team and the following equations and numbers are used to provide a preliminary analysis of the proposed aircraftâ&#x20AC;&#x2122;s dynamics.

During both cruise and loiter, the aircraft has an unstable short period. This instability makes the aircraft uncontrollable by a human pilot because humans are incapable of performing the necessary adjustments at the speeds required to keep the aircraft stable. Longitudinal instability will however increase maneuvering performance of an aircraft which makes it valuable for fighter aircraft, but in this case can lead to structural failure when the aircraft pulls a higher load factor than for what it was designed. Also, typically the short period and phugoid modes are oscillatory in nature, however for this aircraft only the phugoid is oscillatory while the short period response consists of two first order modes, one of which is unstable. đ?&#x2018;˘(đ?&#x2018; ) â&#x2C6;&#x2019;.072728(đ?&#x2018; + 0.19)(đ?&#x2018; 2 â&#x2C6;&#x2019; 4.987đ?&#x2018; + 456) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;5.4345(đ?&#x2018; â&#x2C6;&#x2019; 52.21)(đ?&#x2018; 2 + .006192đ?&#x2018; + .007137) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x2018;&#x17E;(đ?&#x2018; ) 0.40579đ?&#x2018; (đ?&#x2018; + 0.5042)(đ?&#x2018; â&#x2C6;&#x2019; .01591) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x153;&#x192;(đ?&#x2018; ) 0.40579(đ?&#x2018; â&#x2C6;&#x2019; .01591)(đ?&#x2018; + .5042) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x2018;˘(đ?&#x2018; ) 14869đ?&#x2018; (đ?&#x2018; â&#x2C6;&#x2019; 1.489)(đ?&#x2018; + 2.025) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;2070.5561đ?&#x2018; (đ?&#x2018; + 5.894) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) đ?&#x2018;&#x17E;(đ?&#x2018; ) â&#x2C6;&#x2019;17.4689đ?&#x2018; (đ?&#x2018; + 1.051) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301)

61 | P a g e

đ?&#x153;&#x192;(đ?&#x2018; ) â&#x2C6;&#x2019;17.4689(đ?&#x2018; + 1.051) = (đ?&#x2018; ) (đ?&#x2018; đ?&#x203A;żđ?&#x2018;&#x2021; + 2.017)(đ?&#x2018; â&#x2C6;&#x2019; 1.515)(đ?&#x2018; 2 + .03946đ?&#x2018; + .01301) The longitudinal transfer functions for loiter: đ?&#x2018;˘(đ?&#x2018; ) â&#x2C6;&#x2019;.092269(đ?&#x2018; + 0.1783)(đ?&#x2018; 2 â&#x2C6;&#x2019; 1.899đ?&#x2018; + 189) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;4.0842(đ?&#x2018; â&#x2C6;&#x2019; 40.46)(đ?&#x2018; 2 + .004367đ?&#x2018; + .005594) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x2018;&#x17E;(đ?&#x2018; ) 0.3047đ?&#x2018; (đ?&#x2018; + 0.3302)(đ?&#x2018; â&#x2C6;&#x2019; .009055) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x153;&#x192;(đ?&#x2018; ) 0.3047(đ?&#x2018; â&#x2C6;&#x2019; .009055)(đ?&#x2018; + .3302) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x2018;˘(đ?&#x2018; ) 17714đ?&#x2018; (đ?&#x2018; â&#x2C6;&#x2019; 1.155)(đ?&#x2018; + 1.512) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;1665.4722đ?&#x2018; (đ?&#x2018; + .08153) = (đ?&#x2018; ) (đ?&#x2018; đ?&#x203A;żđ?&#x2018;&#x2021; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x2018;&#x17E;(đ?&#x2018; ) â&#x2C6;&#x2019;0.26032đ?&#x2018; (đ?&#x2018; + 20.97) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643) đ?&#x153;&#x192;(đ?&#x2018; ) â&#x2C6;&#x2019;.26032(đ?&#x2018; + 20.97) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + 1.515)(đ?&#x2018; â&#x2C6;&#x2019; 1.16)(đ?&#x2018; 2 + .005814đ?&#x2018; + .005643)

To: Out(2)

Amplitude

In addition to the unstable pole for the short period mode, the elevator responses all have at least one unstable zero and in the case of the flight speed response there is a complex unstable zero. The majority of throttle responses have a zero on the imaginary axis and the flight speed response to the throttle also has an unstable zero. However, because the unstable pole dominates the response, the effects of the unstable zeros is negligible and the response diverges Step Response as can be seen in Figure 9.1.1. The From: In(2) 0 blue line represents the response at cruise and the green ling the response -500 during loiter. During the first second of the response the magnitude grows -1000 from zero to over 3000 during cruise and to slightly under 1000 during -1500 loiter. Though the aircraft is more controllable during loiter it is still -2000 unstable and the aircraft requires control augmentation to be capable of -2500 being flown.

-3000

-3500

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

Time (seconds)

Figure 9.1.1: Vertical Velocity Throttle Response.

1

In the lateral-directional axis the UAS is stable in the Dutch roll, roll, and spiral modes during cruise, but during loiter the spiral mode is unstable, however the Dutch roll and 62 | P a g e

roll modes remain stable. Because the unstable spiral mode has a very low eigenvalue, the aircraft is still controllable in the lateral-directional axis, however the pilot would be required to make adjustments to the aircraft on a periodic basis to remain in control of the aircraft. The lateral transfer functions for cruise are below: đ?&#x203A;˝(đ?&#x2018; ) 0.087599(đ?&#x2018; + 5.744)(đ?&#x2018; + 0.02376) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x2018;?(đ?&#x2018; ) 2.1984đ?&#x2018; (đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x2018;&#x;(đ?&#x2018; ) â&#x2C6;&#x2019;0.092053(đ?&#x2018; + 4.388)(đ?&#x2018; + 0.3463)(đ?&#x2018; â&#x2C6;&#x2019; 0.2703) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x153;&#x2122;(đ?&#x2018; ) 2.1984(đ?&#x2018; 2 + 0.1688đ?&#x2018; + 0.3778) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x203A;˝(đ?&#x2018; ) 0.0055624(đ?&#x2018; + 40.82)(đ?&#x2018; + 2.094)(đ?&#x2018; â&#x2C6;&#x2019; 0.0165) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x2018;?(đ?&#x2018; ) 0.10094đ?&#x2018; (đ?&#x2018; â&#x2C6;&#x2019; 3.609)(đ?&#x2018; + 1.803) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x2018;&#x;(đ?&#x2018; ) 0.22781(đ?&#x2018; + 2.084)(đ?&#x2018; 2 + 0.00157đ?&#x2018; + 0.0631) =â&#x2C6;&#x2019; (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) đ?&#x153;&#x2122;(đ?&#x2018; ) 0.10094(đ?&#x2018; + 1.803)(đ?&#x2018; â&#x2C6;&#x2019; 3.609) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 2.07)(đ?&#x2018; + 0.001854)(đ?&#x2018; 2 + 0.1336đ?&#x2018; + 0.7103) The lateral transfer functions for loiter: đ?&#x203A;˝(đ?&#x2018; ) 0.064931(đ?&#x2018; + 4.473)(đ?&#x2018; + 0.02736) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x2018;?(đ?&#x2018; ) 1.4139đ?&#x2018; (đ?&#x2018; 2 + 0.136đ?&#x2018; + 0.2837) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x2018;&#x;(đ?&#x2018; ) 0.067331(đ?&#x2018; + 3.07)(đ?&#x2018; + 0.3814)(đ?&#x2018; â&#x2C6;&#x2019; 0.2994) =â&#x2C6;&#x2019; (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) đ?&#x153;&#x2122;(đ?&#x2018; ) 1.4139(đ?&#x2018; 2 + 0.136đ?&#x2018; + 0.2837) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x203A;˝(đ?&#x2018; ) 0.0055219(đ?&#x2018; + 31.65)(đ?&#x2018; + 1.346)(đ?&#x2018; â&#x2C6;&#x2019; 0.02311) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x2018;?(đ?&#x2018; ) 0.077716đ?&#x2018; (đ?&#x2018; â&#x2C6;&#x2019; 2.509)(đ?&#x2018; + 1.262) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x2018;&#x;(đ?&#x2018; ) 0.1754(đ?&#x2018; + 1.341)(đ?&#x2018; 2 â&#x2C6;&#x2019; 0.02123đ?&#x2018; + 0.508) =â&#x2C6;&#x2019; (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508) đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) đ?&#x153;&#x2122;(đ?&#x2018; ) 0.077716(đ?&#x2018; + 1.262)(đ?&#x2018; â&#x2C6;&#x2019; 2.509) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 1.336)(đ?&#x2018; â&#x2C6;&#x2019; 0.001365)(đ?&#x2018; 2 + 0.07871đ?&#x2018; + 0.508)

63 | P a g e

Like the longitudinal transfer functions there are unstable zeros in both the aileron and rudder transfer functions. The frequency response of the lateral-directional axis behaves differently than that of the longitudinal axis. The amplitude responses follow each other significantly better than the longitudinal axis, but the loiter low frequency response is consistently 180° out of phase with the cruise frequency response. The phase response converges around 0.01 radian per second and remains close into the higher frequencies. A summary of the mode responses is provided in Table 9.3.2.

9.2) Mil-Std Requirements The MIL-F-8785C requirements for the five modes of dynamic stability must be met by this military UAV. These basic stability requirements mostly involve constraining the damping ratio as well as the natural frequency of the different modes, to prevent damaging vibrations or extreme forces on the aircraft.

The two longitudinal modes, phugoid and short-period, are examined first. The phugoid mode is easily managed by altering the flight of the aircraft, so the sole requirement of a minimum damping ratio is fairly lax.

The short-period mode is more difficult to manage, and therefore has more stringent requirements to affirm stability. There are charts with multiple variables that designate â&#x20AC;&#x153;acceptableâ&#x20AC;? areas for short-period stability, given different flight categories and flight levels. These charts largely address the natural frequency of the short period mode for an aircraft. There is also a short-period damping ratio requirement, which defines the upper and lower bounds for the damping ratio for this mode.

The three lateral modes, namely Dutch roll, roll, and spiral, are then addressed. Dutch roll has the most stringent requirements, since minimum damping ratio, minimum frequency, and minimum product of damping ratio and frequency conditions must be met. Roll mode must be stable and have a time constant that is lower than a maximum value, so that any oscillations are manageable. The spiral mode does not need to be stable, but there is a restriction on the time for the unstable mode to double in amplitude. The following figures and tables should demonstrate that all the basic dynamic stability requirements imposed by the military are met for this UAV.

Figure 9.2.1:Phugoid Requirements for Damping Ratio (from MIL-F-8785C)

10.2.1 Longitudinal Dynamic Stability The above figure (Figure 9.2.1) lists the phugoid damping ratio requirements for the aircraft. Thus far, only the damping ratios for the aircraft during the cruise and loiter segments of the flight path have been calculated, which are both Category B flight conditions. The aircraft is assumed to operate in Level 1 for the duration of the flight, which means the phugoid damping ratio must be greater than 0.04 for both segments.

The actual phugoid damping ratio for the aircraft during cruise is 0.231, and for loiter is 0.0684. Both of these 64 | P a g e

damping ratios are greater than 0.04, so this MIL-F-8785C requirement is met. The augmented stability characteristics, which were improved over the original characteristics, include a phugoid damping ratio of 0.9 for both the cruise and loiter phases, which is well above the minimum specified.

As Figure 9.2.2 demonstrates, the red squares mark the positions on the graph (according to the axes) where the short-period frequencies for this system lie. Since these plotted points are outside of the region bounded by the Level 1 lines, the system operates within Level 2 instead of Level 1. However, the control system was augmented to improve its dynamics stability characteristics. This augmented control system shifted the cruise and loiter markings to within the Level 1 region, which is desired. Therefore, the short-period frequency for both loiter and cruise phases meets the requirements outlined by the MIL-F-8785C.

Figure 9.2.2: Short-period frequency requirements (Category B Flight Phases â&#x20AC;&#x201C; cruise and loiter). Red squares denote the un-augmented shortperiod modes, and orange squares denote the augmented short-period nodes.

65 | P a g e

Figure 9.2.3: Short-period damping ratio requirements (Category B Flight Phases due to loiter and cruise segments being examined)

The short-period damping ratio minimum and maximum limits are given above, in Figure 9.2.3. The aircraft remains in Category B for both cruise and loiter, which were the primary flight patterns considered. The aircraft remains at Level 1 flight conditions, which give the most exacting damping ratio requirements for the short period, as depicted above. The short-period damping ratio for the cruise part of the flight is 0.327, and the short-period damping ratio for the loiter part of the flight is 0.246. Therefore, the loiter part of the flight requires Level 2 conditions, since the Level 1 requirement is not met. However, both the cruise and loiter short-period damping ratios remain within acceptable bounds for at least Level 2 according to MIL-F-8785C. For the augmented design, the short period damping ratio becomes 0.8 for both the cruise and loiter phases, which satisfies the requirements for Level 1, Category B flight at all times.

10.2.2 Lateral Dynamic Stability Figure 9.2.4 tabulates the damping ratio, frequency, and product of damping ratio and frequency requirements for the Dutch roll mode according to MIL-F-8785C. The aircraft remains in Category B for the cruise and loiter phases, so the requirements for Category B at the desired Level 1 will be examined. The minimum Dutch roll damping ratio for these conditions is 0.08 according to Figure 5, and the Dutch roll damping ratios are 0.079 and 0.0552 for cruise and loiter, respectively. Therefore, the un-augmented Dutch roll damping ratio requirement is not met.

66 | P a g e

The second part of the Dutch roll mode requirements that must be examined is the natural frequency.

Since the cruise and loiter flight phases have frequencies of 0.843 rad/s and 0.713 rad/s respectively, the Level 1 minimum frequency requirement of 0.4 rad/s is met on both fronts.

Figure 9.2.4: Dutch Roll damping ratio, frequency, and product of damping ratio and frequency requirements (Category B loiter and cruise Flight Phases, at Level 1 flight conditions)

Additionally, the product of the damping ratio and the frequency must be above a certain minimum, which is 0.15 for Level 1 in Category B flight. However, given damping ratios of 0.079 and 0.0552 and frequencies of 0.843 rad/s and 0.713 rad/s, the resulting products are 0.067 and 0.0394, which are below the requisite value of 0.15. Therefore, the unaugmented Dutch roll product requirement is not met.

Fortunately, the augmented design of the control system gives better results for both the Dutch roll damping ratio and the product of the damping ratio and frequency. The Dutch roll damping ratio becomes 0.4 for both the cruise and loiter flight patterns, which meets the minimum requirement of 0.08. The augmented Dutch roll product for the damping ratio and natural frequency is 0.2 for both cruise and loiter, which exceeds the minimum required value of 0.15. Since the minimum frequency requirement is also met for the cruise and loiter phases (as both are 0.5 rad/s and must be greater than 0.4 rad/s), all the MIL-F-8785C Dutch roll stability requirements are satisfied.

Figure 9.2.6 above contains the MIL-F-8785C basic time constant requirements for roll mode stability.

Since the UAV is a Class II aircraft flying in cruise and loiter flight patterns, both of which are Category B, and the flight conditions are at Level 1, the maximum roll mode time constant is 1.4 seconds. The cruise and loiter roll mode time constants are 0.483 seconds and 0.7463 seconds, respectively, both of which are less than 1.4 seconds. Therefore, the roll mode time constant requirement from MIL-F-8785C is met for both flight phases examined. After augmentation, the roll mode time constant becomes 0.25 for both the cruise and loiter phases, which satisfies the MIL-F-8785C requirement by an even greater margin.

67 | P a g e

The MIL-F8785C standards dictate that the time taken for the amplitude of the spiral mode to double should be more than a prescribed amount, detailed in Figure 9.2.5 above. Since cruise and loiter in Level 1 were considered for this UAV, both of which Figure 9.2.6: Roll mode time constant requirements (Category B Flight are Category B Phases given loiter and cruise) flight patterns, that minimum amount of time is 20 seconds. The time to double (or half, since the cruise spiral mode is stable) amplitude for the spiral mode is 374.7 seconds for cruise and 505.95 seconds for loiter. These values are well above the threshold outlined by the MIL-F-8785C requirements. After augmentation, the spiral mode becomes stable, and therefore never doubles in amplitude, so the MIL-F-8785C condition is met.

Figure 9.2.5: Spiral mode time to double amplitude requirements (Category B Flight Phases of loiter and cruise). The first-order time to double (or half) amplitude is derived from the frequency and damping ratio, which give the spiral mode time constant (Ď&#x201E; = 1/(ÎśĎ&#x2030;))

68 | P a g e

9.3) Augmented Dynamics In order to achieve the level 1 flying qualities enumerated in MIL-F-8785C, a controller has to be implemented to move the poles of the aircraft to locations consistent with the standards. To shift the poles to the desired locations in the complex plane, a simple state feedback controller was implemented of the design in Figure 9.3.1. To calculate K or the gain matrix, the desired poles were determined from MIL-F-8785C using the equation đ?&#x2018; 2 + 2đ?&#x153;&#x201D;đ?&#x2018;&#x203A; đ?&#x153; đ?&#x2018; + đ?&#x153;&#x201D;đ?&#x2018;&#x203A;2 to convert the given natural frequency and damping ratios specified into a form that could be factored for complex roots. In the case of the spiral and roll modes, because they are first order poles the desired pole location did not require conversion into a complex factored form. Upon selecting the desired pole locations, they were placed in a vector (P) and then the MATLAB function place(A,B,P) was used to get the K matrix that multiplies the states for closed loop feedback. A summary of the gain matrices is in Table 9.3.1: Summary of Gains.

Figure 9.3.1: State Feedback Control Block Diagram

Table 9.3.1: Summary of Gains

Elevator Cruise Engine Loiter Elevator Engine Cruise Loiter

Aileron Rudder Aileron Rudder

Î&#x201D;u Î&#x201D;w Î&#x201D;q Î&#x201D;θ -0.00734 0.012084 3.099624 0.427981 3.09E-06 -2.99E-06 0.000742 -0.00172 -0.02457 0.014437 4.748859 0.373785 1.62E-06 -1.42E-06 0.000254 -0.00161 Î&#x201D;β Î&#x201D;p Î&#x201D;r Î&#x201D;Ď&#x2020; -1.35739797 0.975957465 0.481900106 0.181303656 -0.62751626 -0.543853016 -1.103979988 -0.037606106 -1.05137571 2.075763474 0.609300661 0.279591475 -0.26946833 -1.217206809 -1.645262956 -0.041012894

As can be seen from Table 9.3.1: Summary of Gains, the gain is dependent upon the flight speed of the aircraft. This suggests that the overall control scheme that will be implemented will have gain scheduling that modifies the gain as a function of speed and altitude. In addition to gain scheduling, the final control scheme will also include various filters to eliminate structural and engine vibrations from inducing responses from the dynamics of the aircraft. These filters canâ&#x20AC;&#x2122;t be designed at this time without experimental testing of the aircraftâ&#x20AC;&#x2122;s structure and the interaction of the engineâ&#x20AC;&#x2122;s various thrust settings with the structure of the aircraft. Additional filters may be placed on the control inputs as well to facilitate better tracking of the 69 | P a g e

commanded input when the actuator dynamics are also included into the control scheme. Filters may also be placed upon some of the feedback states to shape the response characteristics to various input types so that the aircraft requires less pilot work load.

Table 9.3.2: Summary of Mode Responses

Cruise Frequency Damping

Loiter Frequency Damping

Short Period

2.02

1

1.52

1

1.51

-1

1.16

-1

Phugoid

0.114

0.173

0.0751

Dutch Roll

0.843 2.07

0.0792 1

Spiral

0.00185

1

Augmented Frequency Damping 1

0.8

0.0387

0.1

0.9

0.713 1.34

0.0552 1

0.5 4

0.4 1

0.00137

-1

0.1

1

The longitudinal cruise transfer functions: đ?&#x2018;˘(đ?&#x2018; ) â&#x2C6;&#x2019;.072728(đ?&#x2018; + 57.17)(đ?&#x2018; + 2.827)(đ?&#x2018; + .11) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;5.4345(đ?&#x2018; â&#x2C6;&#x2019; 52.21)(đ?&#x2018; 2 + .05442đ?&#x2018; + .001465) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) đ?&#x2018;&#x17E;(đ?&#x2018; ) 0.40579đ?&#x2018; (đ?&#x2018; + 0.4914)(đ?&#x2018; + .04275) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) đ?&#x153;&#x192;(đ?&#x2018; ) 0.40579(đ?&#x2018; + .04275)(đ?&#x2018; + .4914) = đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x2018;˘(đ?&#x2018; ) 14869đ?&#x2018; (đ?&#x2018; + .07878)(đ?&#x2018; 2 + 1.649đ?&#x2018; + 1.063) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;2663.3802(đ?&#x2018; â&#x2C6;&#x2019; 5.925)(đ?&#x2018; â&#x2C6;&#x2019; .02536) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) đ?&#x2018;&#x17E;(đ?&#x2018; ) 26.7964đ?&#x2018; (đ?&#x2018; + .5332) = 2 đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x153;&#x192;(đ?&#x2018; ) 26.7964(đ?&#x2018; + .5332) = đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x2018;˘(đ?&#x2018; ) â&#x2C6;&#x2019;.092269(đ?&#x2018; + .07185)(đ?&#x2018; 2 + 14.17đ?&#x2018; + 53.12) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;4.0842(đ?&#x2018; â&#x2C6;&#x2019; 40.46)(đ?&#x2018; 2 + .03367đ?&#x2018; + .0006336) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; ) đ?&#x2018;&#x17E;(đ?&#x2018; ) 0.3047đ?&#x2018; (đ?&#x2018; + 0.3237)(đ?&#x2018; + .02611) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2019; (đ?&#x2018; )

70 | P a g e

đ?&#x153;&#x192;(đ?&#x2018; ) 0.3047(đ?&#x2018; + .02611)(đ?&#x2018; + .3237) = 2 (đ?&#x2018; ) (đ?&#x2018; đ?&#x203A;żđ?&#x2018;&#x2019; + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x2018;˘(đ?&#x2018; ) 17714đ?&#x2018; (đ?&#x2018; + .0312)(đ?&#x2018; 2 + 1.714đ?&#x2018; + 1.157) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) đ?&#x2018;¤(đ?&#x2018; ) â&#x2C6;&#x2019;3442.7169(đ?&#x2018; â&#x2C6;&#x2019; 20.14)(đ?&#x2018; â&#x2C6;&#x2019; .002741) = (đ?&#x2018; 2 + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) đ?&#x2018;&#x17E;(đ?&#x2018; ) 132.3281đ?&#x2018; (đ?&#x2018; + .3318) = 2 (đ?&#x2018; ) (đ?&#x2018; đ?&#x203A;żđ?&#x2018;&#x2021; + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) đ?&#x153;&#x192;(đ?&#x2018; ) 26.7964(đ?&#x2018; + .5332) = 2 đ?&#x203A;żđ?&#x2018;&#x2021; (đ?&#x2018; ) (đ?&#x2018; + .18đ?&#x2018; + .01)(đ?&#x2018; 2 + 1.6đ?&#x2018; + 1) Utilizing state feedback control to augment the dynamics of the aircraft turned the unstable and uncontrollable short period mode into a stable mode. The pair of single order poles were combined into a stable complex pair similar to the phugoid mode. Because of the pole placement used to select the gains, the short period and phugoid mode settles quickly without any transient oscillations. Another benefit of the augmentation of the aircraft dynamics was the elimination of unstable zeros for all the transfer functions except those that model the z-axis velocity response. The z-axis velocity response to elevator inputs kept its one unstable zero and the complex pair of zeros shifted closer to the imaginary axis. The z-axis velocity response to an increase in throttle however shifted a zero that was on the imaginary axis into the right half plane making that zero unstable. These unstable zeros result in a large initial overshoot which then transitions Impulse Response From: In(2)

4

5

x 10

4

To: Out(2)

Amplitude

3

2

1

0

-1

-2

0

10

20

30

40

50

60

70

80

90

100

Time (seconds)

Figure 9.3.2: Z-axis Velocity Throttle Impulse Response

into undershoot before reaching steady state as can be seen in Figure 9.3.2, the blue line represents the response during cruise and the green line is the response during loiter. Additional filters will be needed to smooth out the response and improve input tracking for the vertical velocity responses. The responses from the other longitudinal transfer functions settle to their steady state responses in a similar way with much less or no undershoot when subject to an impulse input. The response to a step input in two cases responds with 71 | P a g e

overshoot and then undershoot for the z-axis velocity and pitch rate responses to elevator input. In the remainder of the cases however, the system settles to steady state with either overshoot or undershoot except for the pitch angle response to throttle which settles directly to steady state. An example step response is shown in Figure 9.3.3, the blue line represents the cruise flight condition and the green line represents the loiter flight condition. Graphical comparison of the un-augmented response with the augmented response was not possible because of the rate at which the un-augmented system diverges prevents details about the Step Response From: In(2)

4

14

x 10

12

Amplitude

To: Out(1)

10

8

6

4

2

0

0

10

20

30

40

50

60

70

80

90

100

Time (seconds)

Figure 9.3.3: Flight Speed Throttle Response

augmented system from being seen.

The augmented lateral-directional cruise transfer functions: đ?&#x203A;˝(đ?&#x2018; ) 0.093684(đ?&#x2018; + 8.138)(đ?&#x2018; + 0.07268) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25) đ?&#x2018;?(đ?&#x2018; ) 2.1984đ?&#x2018; (đ?&#x2018; 2 + .4121đ?&#x2018; + 0.2479) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25) đ?&#x2018;&#x;(đ?&#x2018; ) â&#x2C6;&#x2019;0.092053(đ?&#x2018; + 7.286)(đ?&#x2018; + 0.228)(đ?&#x2018; â&#x2C6;&#x2019; 0.1493) = (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25) đ?&#x203A;żđ?&#x2018;&#x17D; đ?&#x153;&#x2122;(đ?&#x2018; ) 2.1984(đ?&#x2018; 2 + 0.4121đ?&#x2018; + 0.2479) = đ?&#x203A;żđ?&#x2018;&#x17D; (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25)

72 | P a g e

𝛽(𝑠) 0.0055624(𝑠 + 40.82)(𝑠 + 4.08)(𝑠 + 0.09936) = (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝛿𝑟 (𝑠) 𝑝(𝑠) 0.10094𝑠(𝑠 2 + 0.7049𝑠 + 0.1922) = 𝛿𝑟 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝑟(𝑠) 0.22781(𝑠 + 4.057)(𝑠 + 0.1092)(𝑠 + 0.02895) =− (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝛿𝑟 (𝑠) 𝜙(𝑠) 2.1984(𝑠 2 + 0.4121𝑠 + 0.2479) = 𝛿𝑟 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) The augmented lateral-directional loiter transfer functions: 𝛽(𝑠) 0.073822(𝑠 + 7.854)(𝑠 + 0.07176) = 𝛿𝑎 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝑝(𝑠) 1.4139𝑠(𝑠 2 + .417𝑠 + 0.2467) = 𝛿𝑎 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝑟(𝑠) 0.067331(𝑠 + 7.478)(𝑠 + 0.2294)(𝑠 − 0.1679) =− (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝛿𝑎 (𝑠)

Impulse Response From: In(1) 0.9

0.8

0.7

0.6

To: Out(4)

Amplitude

0.5

0.4

0.3

0.2

0.1

0

0

5

10

15

20

25

30

35

40

Time (seconds)

Figure 9.3.4: Roll Angle Aileron Impulse Response

𝜙(𝑠) 1.4139(𝑠 2 + 0.417𝑠 + 0.2467) = 𝛿𝑎 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝛽(𝑠) 0.0055219(𝑠 + 31.67)(𝑠 + 4.093)(𝑠 + 0.09936) = (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 𝛿𝑟 (𝑠) 𝑝(𝑠) 0.077716𝑠(𝑠 2 + 0.7628𝑠 + 0.174) = 𝛿𝑟 (𝑠) (𝑠 + 4)(𝑠 + 0.1)(𝑠 2 + 0.4𝑠 + 0.25) 73 | P a g e

đ?&#x2018;&#x;(đ?&#x2018; ) 0.1754(đ?&#x2018; + 4.067)(đ?&#x2018; + 0.1085)(đ?&#x2018; + 0.01938) =â&#x2C6;&#x2019; (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25) đ?&#x203A;żđ?&#x2018;&#x; đ?&#x153;&#x2122;(đ?&#x2018; ) 0.077716(đ?&#x2018; 2 + 0.7628đ?&#x2018; + 0.174) = đ?&#x203A;żđ?&#x2018;&#x; (đ?&#x2018; ) (đ?&#x2018; + 4)(đ?&#x2018; + 0.1)(đ?&#x2018; 2 + 0.4đ?&#x2018; + 0.25) Impulse Response From: In(1) 0.2

0.15

0.1

To: Out(3)

Amplitude

0.05

0

-0.05

-0.1

-0.15

0

5

10

15

20

25

30

35

40

Time (seconds)

Figure 9.3.5: Yaw Rate Aileron Impulse Response

Augmentation of the lateral-directional dynamics also had a significant effect upon the dynamics of the aircraft. Except the yaw rate response to the aileron, every transfer function that had unstable zeros before augmentation, did not have them after augmentation. Also the yaw rate response to the rudder after augmentation transformed a complex pair of zeros into separate stable zeros, improving the response characteristics. Like the longitudinal response, the lateral-directional response, when augmented, settles much faster and has less of an overshoot than the un-augmented dynamics. The exception to this is the sideslip response to rudder inputs which has a higher overshoot, but it still settles significantly faster than the unaugmented dynamics. The roll angle and roll rate responses, when augmented, settle to steady state without any oscillations as can be seen in Figure 9.3.4. The red and blue lines are the un-augmented dynamics for cruise and loiter respectively, and the green and teal lines are the augmented dynamics for those same flight conditions. The yaw rate and sideslip angle responses are not as damped as the roll angle and rate responses, and have additional oscillations before the transients die down as can be seen in Figure 9.3.5.

10) Weight Justification 10.1) Detailed Weight Analysis The aircraft's weight was initially estimated based on the mission design requirements and geometry comparable to similar aircrafts. Then a detailed weight analysis was performed which collaborates the methods established by the General Dynamics, Torenbeek [35], and Vought [36]. The detailed weight analysis (Class II [36]) was utilized to precisely project the weight of the airplane using established methods that are based on the geometry of aircraft. However, these methods are overly general and not sufficiently accurate when applied to Ancile. These methods are less precise when applied to Ancile due to its 74 | P a g e

unconventional characteristics; such as electric laser including turret, high aspect ratios, cooling system, high capacity energy storage, and single aft engine. Therefore, a more in-depth analysis, outlined in the following section, was conducted after these initial estimates in order to determine the total weight.

Table 10.1.1 presents the Class II detailed empty weight estimation using the General Dynamics, Torenbeek, and Vought methods.

Table 10.1.1 Summary of Class II Weight Estimation

19827.33

Torenbeek Method (lbs.)

28545.70

Vought Method (lbs.)

33071.95

Average Values (lbs.)

27148.32

956.58

2530.33

1120.61

1535.84

766.65

1251.47

1049.61

1022.58

10215.48

----

19376.26

14795.87

585.09

571.12

370.08

508.76

Weight Component

GD Method (lbs.)

Wing Vertical Tail Canard Fuselage Nose Landing Gear Main Landing Gear

4680.39

4568.65

3138.65

4129.23

Structure

37031.53

37467.27

58127.16

49140.60

Engines Fuel System

----

13046.41

----

13046.41

----

742.58

217.81

480.20

Power Plant

----

13788.99

217.81

13526.61

Flight Control System Instruments/Avionics/Electronics Auxiliary Power Unit Hydraulic and Pneumatic System Electrical System

2596.16

----

1418.24

2007.20

2489.70

1764.85

2414.90

2223.15

----

1022.01

3300.00

2161.00

----

1226.41

416.61

821.51

1923.63

----

1037.35

1480.49

Fixed Equipment

7009.49

4013.26

8587.10

10119.65

Empty Weight

----

----

----

72786.86

Maximum Takeoff Weight

----

----

----

230582.30

10.2) Class III Weight Estimation The aircraft's total weight was determined piece-wise according to Class III weight analysis by taking into account the materials of each individual component, which were dependent upon fiscal and structural necessity. Although tedious, the component-by-component approach allowed for a more accurate estimation of the total weight. The resulting component-by-component estimations are displayed in exploded weight groups within the foldout section at the end of this section.

10.3) Center of Gravity Analysis The total center of gravity of the aircraft was determined by finding the local center of gravity of each aircraft component and using this in conjunction with the component’s position relative to the aircraft to determine the global cg of the component. The component cg’s were then summed to determine the total aircraft center of gravity. The locations of each component’s center of gravity are displayed below in Table 10.3.1.

Table 10.3.1: Detailed CG location and Moments of individual aircraft components

Component

Weight (lbs)

x-cg (ft)

x-Moment (ft-lb)

z-cg (ft)

z-Moment (ft-lb) 75 | P a g e

Wing

27148.3

1171.07

31792589

-72.48

-1967710

Fuselage

14795.9

987.62

14612699

2.78

41132

Empennage

2558.4

1250.54

3199389

112.23

287133

Nose Gear

508.8

308.01

156703

-167.77

-85354

Main Gear

4129.2

1242.17

5129168

-183.03

-755767

Propulsion

17095.2

1646.98

28155390

-0.04

-732.5

Electrical

5870.2

1547.94

9086697

-0.05

-283.5

Avionics

400.0

1710.00

684000

18.04

7216.0

Cooling system

3522.9

726.00

2557625

1.78

6270.8

Turret

5500.0

0.00

0

0.00

0.0

Power System

10732.7

880.84

9453870

2.07

22248.8

Laser

11025.0

700.10

7718603

-0.20

-2205.0

Fuel

114913.2

1118.24

128500520

-59.52

-6839632

The following table displays the resulting mass and the x and z coordinates of the center of gravity for both the TOGW and the operating empty weight conditions.

Table 10.3.2: CG Shift under Various Loading Conditions

Fueled/Un-fueled

Weight (lbs)

x-cg (ft)

y-cg (ft)

z-cg (ft)

at OEW

104090

87.6

0

-7.2

at TOGW

219003

90.4

0

-6.0

The fuel burn results in shift of the center of gravity equal to 6.9 % of the mean geometric chord as the aircraft goes from initial takeoff weight to empty weight. Table 10.3.3 and Table 10.3.4 present a detailed takeoff weight summary and moments of inertia, respectively.

Figure 10.3.1: Location of Various Component CGs

76 | P a g e

.

Table 10.3.3 Detailed Takeoff Weight Wfix

30,780.65 lb

Wstructure

49,140.56 lb

WPP

22,965.38 lb

WE

103,286.58 lb

WTO

218,199.77 lb

Table 10.3.4: Moments of Inertia đ??źđ?&#x2018;Ľđ?&#x2018;Ľ đ??ľ

4,115,323.92 slugs-ft2

đ??źđ?&#x2018;Śđ?&#x2018;Ś đ??ľ

5,273,981.75 slugs-ft2

đ??źđ?&#x2018;§đ?&#x2018;§ đ??ľ

11,022,367.06 slugs-ft2

đ??źđ?&#x2018;Ľđ?&#x2018;§ đ??ľ

80,000 slugs-ft2

11) Performance 11.1) Takeoff Performance The required takeoff field length was determined using the relations presented in Fundamentals of Aircraft and Airship Design, Chapter 10. The flaps were used according to the configuration specified in 4.2), with a deflection of 12 degrees for takeoff, according to the suggested value for fowler flaps. Assuming twelve degree deflection of the flaps during takeoff the lift coefficient attainable is CLmax=2.2. The coefficient of rolling friction for a concrete surface was assumed to be approximately 0.025 (based on the average value outlined in Fundamentals of Aircraft Design, ch10, Tble 10.3 pg 261) [37]. The trajectory is displayed below.

Figure 11.1.1: Takeoff Flight Path

77 | P a g e

Figure 11.1.2: ROC vs. velocity at various altitudes including cruise altitude, operational ceiling, and ceiling.

Table 11.1.1: Takeoff Parameters

Takeoff Performance

Takeoff Conditions

Ground Distance

5210 ft

Rotation Distance

424 ft

Transition Distance Total Takeoff Distance Takeoff Speed

1366 ft 7000 ft 126 knots

CLmax TO

2.2

CD TO

0.065

L/DTO Î TO

12.4 0.51

Because the ending transition altitude is greater than the height of the 50 ft obstacle the clearance portion of takeoff is unnecessary and thus the aircraft immediately begins its climb without including a clearance portion of takeoff. The throttle setting is low because the entire length of the runway (7000 ft requirement) is being used, at a higher throttle setting the aircraft could take off a shorter runway but this was deemed unnecessary given the design constraints.

11.2) Climb Performance During climbing flight the aircraft flies at an airspeed of 250 knots because, as Figure 11.1.2 shows, this airspeed yields the greatest rate of climb for the aircraft. The climb analysis was done using the maximum continuous thrust that can be supplied by the aircraft at sea level; 60,400 lbf (maximum thrust is 68,600 lbf but this can be maintained for less than 5 minutes). The aircraft climbs with an angle of attack of three degrees and an initial takeoff weight of 204,000 lbs.

The initial maximum rate of climb is 4940 ft/min and the rate of climb at cruise (40,000ft) is 1200 ft/min. The operational ceiling is defined as the altitude where ROC is equal to 300 ft/min for military aircraft. This occurs 78 | P a g e

at an altitude of 65,000 ft, while the absolute ceiling (the altitude at which the aircraft is unable to climb) occurs at an altitude of 150,000 ft.

11.3) Loiter Performance During the loiter phase the aircraft must travel in a trajectory that makes a circuit so that it can target and destroy the ICBM. The difficulty in a circuitous trajectory arises from the fact the laser can only fire within 120 degrees of the aircraft heading. Thus there will be periods during loiter where the aircraft cannot fire on the target. An ideal trajectory will minimize these periods so that the aircraft can fire at the target for a majority of the loiter phase. To determine an optimal configuration, an elliptical (racetrack) configuration was compared to a figure 8 configuration for similar length scales that would put the aircraft in range of the target for the entirety of the circuit. The vision angles were determined for each type of circuit with three different target locations. The target was placed at the center of the circuit, the outside (along y-axis), and inside the circuit (along the x-axis) as shown in figures 12.3.1 and 12.3.2 (all dimensions displayed in nautical miles).

Each configuration is arranged so that the greatest distance between the aircraft and the target is less than or equal to 250 miles, which is the maximum range that the laser can still be used effectively from. Also, since the aircraft can handle a turning radius as small as 0.6 miles turning flight will not be a design constraint.

Figure 11.3.1: Target Locations for Elliptical Flight Pattern

Figure 11.3.2: Target Locations for Figure 8

Both trajectories were analyzed for each target location (6 total configurations) to determine which had the greatest length of time where the target could be fired upon. The results are displayed in Table 11.3.1: Summary of Loiter Pattern Analysis.

Table 11.3.1: Summary of Loiter Pattern Analysis Trajectory Elliptical (Race Track)

Figure Eight (Dual Loops)

Target Location 1. Origin 2. Along y- axis 3. Along x-axis 1. Origin 2. Along y- axis 3. Along x-axis

% of time unable to fire 34 23 34 33 33 16.7

79 | P a g e

Diagrams of the two best configurations are displayed in Figure 11.3.1 and Figure 11.3.2 which denote the regions in which the vision angles are insufficient.

Figure 11.3.1: Elliptical configuration with target outside flight path. Red line denotes portion of flight path where vision angles are insufficient.

Figure 11.3.2: Figure eight configuration with target located along the x axis inside the flight path. Red line denotes portion of flight path where vision angles are insufficient.

The figure eight configuration with the target positioned along the x-axis approximately 50 miles from the center of the circuit was chosen because it provided the maximum percentage of time when the laser could be fired; 83%. Thus for 20 hours of the 24 hour loiter period, the aircraft will be able to destroy the target.

However there is still the possibility of missile launch during the portion of the trajectory in which the vision angles are insufficient. Should this occur the aircraft will pull a 2-g turn (which is 67% of the maximum allowable g-loading; Factor of Safety=1.5) at the loiter velocity, which results in a turning radius of 1.0 mile.

The aircraft would turn at this rate until the vision angles were again adequate and then continue gradually 80 | P a g e

turning until it could rejoin the loiter circuit. The maximum time that this would require would occur at the point where the aircraft heading is exactly 180 degrees with respect to the target. In this case the aircraft must travel 2.6 miles in turning flight to regain vision of the target. At the cruise speed of 321.3 knots this could be completed in 25 seconds. Taking into effect pilot reaction time and transition to turning flight the total time to come about into range of the target can be estimated to be approximately 40 seconds at the point where the required turning angle is at a maximum.

During the 24 hour loiter period the aircraft operates with a lift to drag ratio of 25.7 and thrust specific fuel consumption of 0.534 lbs/(lb-hr). Flying at the cruise speed of 321.27 knots results in a fuel burn of 71,610 lbs over the course of the loiter phase. During the time the aircraft travels a total distance of 8873 miles, which corresponds to 7 circuits of the figure eight loiter pattern.

11.4) Landing Performance The Landing Analysis was performed according to the procedure laid out in Fundamentals of Aircraft Design [37]. The Landing distance was estimated assuming a Maximum Landing Weight (MLW) of 115,604.8 lbs. The MLW is defined by RFP as the total weight of the aircraft with only the reserve fuel. That is, it is maximum empty weight (58,828.1 lbs) plus the reserve fuel weight (5,453 lbs). The coefficient of friction for the concrete runway was assumed to be approximately 0.3 for brakes applied, according to Table 10.3 in Fundamentals of Aircraft Design [37]. The aircraft must clear a fifty foot obstacle and be able to land and slow to a stop before the end of the runway (7000 ft); the resulting trajectory is displayed below.

Figure 11.4.1: Landing Flight Path Table 11.4.1: Landing Performance Parameters

Landing Performance Air Distance 2812 ft Free Roll Distance 449 ft Braking Distance 1561 ft Total Landing Distance 4822 ft

Landing Conditions CL max ldg 2.7 CD landing 0.138

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Speed at Obstacle Touchdown Speed

116 knots 89 knots

12) Recommendations 12.1) Design Recommendations 12.1.1)

Aerodynamic and Configuration Recommendations

The aircraft proposed in this proposal could be significantly improved through the use of industry practices that we were unable to utilize due to limited resources and personnel. If we had the resources of a company that was actively engaged in designing and producing aircraft for the Department of Defense there are various design considerations that could be studied in further depth to provide an aircraft that has even better performance than what is currently proposed. The main topic of further investigation would that of the wing’s design. Through the extensive use of computational fluid dynamics and wind tunnel testing to verify the results of computer simulation, we could have designed a higher performing wing planform. Also, investigation into wing tip devices to limit or eliminate wingtip vortices would have been beneficial to having a comprehensive understanding of the optimal wing design for the specific mission. Along the same lines as optimizing the wing design for loiter performance, engaging in computational fluid dynamics analysis of the fuselage to refine the shape to make it as “slippery” as possible as well as reduce interference with the wing and empennage would have been conducive to providing the best possible design proposal.

Another way the design could have been improved is through the design of a custom engine to more closely fit the design needs of the rest of the airframe. There was a conscious decision throughout the design process to utilize technology that is available today to reduce design costs and difficult as well as allow the introduction into service at or before 2035. Designing a custom engine would not have guaranteed the timely introduction into service like selecting a pre-existing engine would. As engine technology improves, the airframe will most likely receive a retrofit or go through a major block change to improve the performance of the UAS during the lifetime of the aircraft, thus the aircraft that is seen toward the end of its service would only be similar to the initial production model at the most superficial level.

13.1.2 Avionics Recommendations 2035 is the earliest expected date of operation of this aircraft, so allowances must be made for improvements in technology throughout the interceding duration. Avionics in particular may change tremendously from their current form to something that approaches science fiction, due to improvements in electronics and processing power following Moore’s Law and other related laws governing networks, data transmission, fiber optics, and computer memory. While the future cannot be accurately predicted, improvements in the range, efficiency, cost, weight, reliability, and durability of avionics are almost assured over the next twenty years. As an example, the flight computers should become cheaper and more powerful, since they would have approximately 222/2 = 2048 as many transistors as the computers currently do (provided Moore’s Law holds true until 2035). Therefore, the avionics system that is finally implemented as part of the aircraft may be markedly different from the designed avionics systems proposed here.

The physical structures, or racks, containing the avionics for this aircraft must also allow for an expanded avionics load out. Therefore, the racks must contain empty spaces or be modular in design, so that additional avionics components may be added in the future. Modular avionics racks seem to be the best design 82 | P a g e

decision, since they may be attached to one another provided space is available, allowing for additional avionics capabilities or at least increased redundancy. If space for additional avionics is limited, then the racks could be designed so that the shelves and inner compartments are adjustable and can be reconfigured as necessary to fit different components. These structures would then require the same volume within the aircraft, but the avionics loud out could be completely different. This idea would reduce high-level design complexity and minimize the use of resources to accommodate different avionics configurations.

Finally, certain components may be replaced by upgraded systems or otherwise made obsolete. For example, the KU-band satellite receiver will likely be replaced by Ka-band satellite receivers by 2035 when the aircraft will begin operation. The Ka-band satellite receivers have the advantage of higher upload and download rates (10 to 20 Mbps) compared to the KU-band satellite receivers (1 to 2 Mbps), but there is currently limited coverage on the Ka-band compared to the KU-band (George 2). While this is only one example of a system’s components becoming obsolete and replaced by upgraded components with similar functions, doubtlessly more systems will undergo the same process. Therefore, some avionics components may become obsolete, consolidated, or otherwise replaced by 2035, meaning that the current proposed avionics system is only preliminary and still a work in progress.

12.2) Capability Upgrades 12.2.1)

Avionics Upgrades

Professor Li of Beihang University suggests that a logical future expansion of avionics would use “cloud” technology [37]. His overall idea is that avionics contained within the aircraft would interface with and interpret signals from other electronics and sensor systems, which may be ground-, air-, or space-based. These systems could contain vital information such as the position of enemies, the terrain and weather conditions beyond the scope of the UAV’s radar, or the nearest friendly airfield. The UAV could also be remotely controlled by computer processors not located on the UAV itself, although a small lag time would be introduced between the transmission of a command and its execution. Similarly, some information processing tasks could be performed at sites with large, immobile computers by quickly relaying the information gathered by the UAV to those sites. These sorts of processing tasks could include image recognition, information storage (including maps, enemy locations, flight data, etc.), trajectory or flight path plotting, redundant resource management, and redundant status monitoring (so that a third-party computer tracks the status of all electronics and avionics systems on the aircraft). This management redundancy could allow for the UAV to continue operating if certain systems are destroyed, such as the flight or mission computers. If the UAV remains connected to the cloud after these systems are destroyed, then an idle computer or computer cluster also within the cloud could possibly perform the flight or mission computer functions remotely, albeit at a loss of speed and data efficiency. Therefore, an avionics cloud would allow the aircraft to remain controllable or recoverable in case of an emergency, and can absorb some of the aircraft computers’ workload. This “cloud” feature of the avionics system will likely become more important as cloud-computing technology and infrastructure develop over the next twenty years.

12.2.2)

Laser System Upgrades

As battery and super-capacitor technology improves, it will become possible to include more power storage on the aircraft, while maintaining the same weight. Due to the architecture of the laser power system, the more laser bursts the aircraft is capable of, the faster the laser can fire. A limiting factor in the increase in fire rate is the current rating of the super-capacitors. If a super-capacitor receives too much current it can overheat or explode rendering the aircraft unable to complete its primary mission. As the super-capacitor technology matures, it can be expected that the current that can be supplied will also increase, reducing charging times for each of the individual super-capacitors. The cooling system that has been designed for this 83 | P a g e

aircraft is capable of sustaining a higher rate of fire than the current power technology allows, thus retrofitting an upgraded system should consist of matching the battery and capacitor capabilities and implementing the architecture on a slightly larger scale.

13) Cost The success of any defense program is determined not only by its performance, but how much the program costs to develop and run. To estimate this important metric, the cost estimation methodology presented in Roskamâ&#x20AC;&#x2122;s Aircraft Design Part VIII was utilized to estimate the costs of the research and development phase in which prototypes for flight trials and assessment are produced as well as to produce the production version of the aircraft and operate it over its minimum 20 year lifespan. The summary of these costs are presented in Table 12.2.1. The cost of developing the prototypes is constant regardless of the number of production aircraft because the same number of prototypes will be produced and the engineering and manufacturing workload is estimated to account for that limited production run. Acquisition and operating costs are dependent on the amount of airframes that are produced during the production run. Numbers are presented for 20, 30, 40 systems to show the per aircraft savings incurred by purchasing a larger number of systems. All dollar amounts are in 2013 US dollars because future monetary trends cannot be accurately predicted for decades in advance.

Table 12.2.1: Summary of Costs

Aircraft Produced Cost Research and Development Acquisition Operating

60 $1,846,626,279.48

90 $1,846,626,279.48

120 $1,846,626,279.48

$4,919,057,187.99 $11,576,754,735.76

$6,673,333,744.59 $17,365,132,103.64

$8,309,341,393.69 $23,153,509,471.52

Total Cost per Aircraft

$18,342,438,203.24 $305,707,303.39

$25,885,092,127.71 $287,612,134.75

$33,309,477,144.69 $277,578,976.21

The individual costs associated with producing the prototype aircraft for testing and evaluation purposes are presented in Table 12.2.2. The airframe engineering cost consists of the costs of paying the engineers during the entire design phase of the prototype aircraft. The cost will vary somewhat depending on the cost of living where the aircraft is being designed. The value was calculated using Roskamâ&#x20AC;&#x2122;s equation for estimating the man hours required to develop the aircraft and multiplying the required man hours by an inflation adjusted rate presented in Roskam as well. The development support costs are the associated modeling and testing costs incurred during the development stage such as wind tunnel and structural tests.

This number is also subject to variations depending on the development test facilities required and the costs to operate those facilities. The flight test airframe cost is the cost of producing eight prototype aircraft for flight testing and certification of various subsystems necessary for the operation of the UAS. The financial outlay required to finance those various activities to ensure the aircraft meets the requirements is enumerated by the flight test operations and it includes the all the expenditures related to conducting flight tests of the prototype aircraft. The profit and financing costs are additional costs associated with developing any aircraft. The cost

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of the aircraft cannot be expected to be paid in one lump sum and so the financing cost will depend upon the interest rates available to the company that is producing the prototypes for evaluation.

Table 12.2.2: Research and Development Cost Breakdown

Research and Development Airframe Engineering Development Support Flight Test Airframes Flight Test Operations Profit Financing Total

$171,750,065.99 $37,325,995.52 $1,282,181,818.98 $34,879,540.57 $152,613,742.11 $167,875,116.32 $1,846,626,279.48

Once the testing and evaluation phase is complete and the UAS is selected to go into full production, the cost of acquiring the UAS is dependent upon the number of airframes to be manufactured. As the number of airframes increases the total cost of the entire program increases, but the cost per unit decreases. Table 12.2.3 enumerates the cost of re-engineering the aircraft after flight testing, producing the production version of the aircraft, and the lifetime operating costs. By doubling the order of aircraft, the cost per aircraft drops almost $12 million and the production program as a whole costs approximately $3.5 billion more. Those additional aircraft, however, cost an additional $550 million dollars to operate over the 20 year lifespan of the aircraft for a total program cost of $33.3 billion. The cost of 20 systems over their 20 year lifetime would be $18.3 billion dollars. The cost per system ranges between $833 and $917 million over the lifetime. If additional systems are purchased by the US Air Force or if the export of the UAS is permitted by the US government, than costs per airframe could be expected to decrease even more. If the UAS is slated to be in service for a significantly longer period than 20 years, than the cost per airframe will increase in order to retrofit upgraded systems into the airframe to make it capable of operating in the future environment. In order to retrofit the UAS significant additional engineering work will have to be undertaken and could potentially increase the per airframe cost by a large margin.

Table 12.2.3: Acquisition and Operating Cost Breakdown

Acquisition Airframes Produced Engineering Cost Avionics and Engine Cost Program Cost Materials Cost Tooling Cost Quality Control Cost Test Operations Cost Financing Profit Total Cost Airframe Cost

60 90 120 $76,582,184.21 $95,709,308.26 $110,167,159.75 $1,358,321,219.76 $2,037,481,829.64 $2,716,642,439.51 $1,561,834,477.17 $1,931,553,871.84 $2,245,818,852.98 $556,815,688.89 $821,293,862.07 $1,067,688,661.51 $250,416,918.19 $297,732,895.16 $333,436,493.45 $203,038,482.03 $251,102,003.34 $291,956,450.89 $35,011,529.33 $52,517,293.99 $70,023,058.66 $406,533,651.90 $551,515,185.50 $686,722,429.23 $447,187,017.09 $606,666,704.05 $755,394,672.15 $4,919,057,187.99 $6,673,333,744.59 $8,309,341,393.69 $81,984,286.47 $74,148,152.72 $69,244,511.61 85 | P a g e

Fuel and Oil Personnel (Direct) Personnel (Indirect) Spares Depot Misc Cost/hr

Operations $4,470,896,651.24 $6,706,344,976.86 $170,100,000.00 $255,150,000.00 $3,348,482,117.40 $5,022,723,176.10 $1,198,421,815.30 $1,797,632,722.94 $1,837,580,116.79 $2,756,370,175.18 $551,274,035.04 $826,911,052.55 $7,294.07

$7,294.07

$8,941,793,302.48 $340,200,000.00 $6,696,964,234.80 $2,396,843,630.59 $3,675,160,233.57 $1,102,548,070.07 $7,294.07

15.0 Software Used Description 15.1 Advanced Aircraft Analysis (AAA) Advanced Aircraft Analysis (AAA) is the industry standard aircraft design, stability, and control analysis software. AAA is installed in over 55 countries and is used by major aeronautical engineering universities, aircraft manufacturers, and military organizations worldwide.

Advanced Aircraft Analysis provides a powerful framework to support the iterative and non-unique process of aircraft preliminary design. The AAA program allows students and preliminary design engineers to take an aircraft configuration from early weight sizing through open loop and closed loop dynamic stability and sensitivity analysis, while working within regulatory and cost constraints.

The design methodology used in Advanced Aircraft Analysis is based on Airplane Design I-VIII, Airplane Flight Dynamics and Automatic Flight Controls, Parts I and II, by Dr. Jan Roskam, and Airplane Aerodynamics and Performance, by Dr. C.T. Lan and Dr. Jan Roskam. AAA incorporates the methods, statistical databases, formulas and relevant illustrations and drawings from these references.

15.2 Star CCM+ v8 STAR-CCM+ provides the world's most comprehensive engineering simulation inside a single integrated package. Much more than just a CFD solver, STAR-CCM+ is an entire engineering process for solving problems involving flow (of fluids or solids), heat transfer and stress.

15.3 GasTurb 11 Gasturb is a gas turbine performance simulation program for Microsoft Windows written and distributed by former MTU Aero Engines employee Joachim Kurzke. Based around a comprehensive set of pre-defined engine configurations, the program performs both design point and off-design performance modeling, parametric studies, cycle optimization and Monte Carlo simulations. The program is widely used both in the power generation and propulsion industries and in teaching institutions.

15.4 Conceptual Flutter Analysis of Cantilever Wing (CFACW) CFACW is a graphical user interface (GUI) flutter prediction code based on the geometric and physical inputs implemented in MATLAB. It helps to give preliminary flutter boundaries during the conceptual design phase. The NASA based CFACW software has been used by researchers and students alike to determine flutter.

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15.5 DesignFOIL DesignFOIL is a Windows-based airfoil software tool. It helps to create, modify, and aerodynamically analyze airfoil shapes. It also performs basic wing layout, CAD export, and creates CFD preparation files.

Although built for professionals, the user-friendly interface is used by many hobbyists as well.

15.6 Ansys CFX ANSYS CFX software is a high-performance, general purpose fluid dynamics program that has been applied to solve wide-ranging fluid flow problems for over 20 years. At the heart of ANSYS CFX is its advanced solver technology, the key to achieving reliable and accurate solutions quickly and robustly. The modern, highly parallelized solver is the foundation for an abundant choice of physical models to capture virtually any type of phenomena related to fluid flow. The solver and its many physical models are wrapped in a modern, intuitive, and flexible GUI and user environment, with extensive capabilities for customization and automation using session files, scripting and a powerful expression language.

15.7 Solidworks Solidworks is a general, interactive 3D-modeling program used by industry and students alike. By using Solidworks, SCAD was able to design the aircraft. Solidworks also has the ability to perform FEA analysis in fluids and thermodynamics, enabling the laser and cooling systems to be designed accurately.

15.8 Rhinoceros w/ VRAY Rhinoceros is a general, interactive 3D-modeling program that can handle even the most complicated model. Rhinoceros is equipped with the VRAY plug in that allows for detailed rendering of the model with various materials and lighting conditions. Using VRAY, SCAD was able to generate all its appropriate drawings.

15.9 Microsoft Excel Excel is a spreadsheet program that is a part of the Microsoft Office suite. Excel was used to create tables and graphs for this proposal as well as perform a myriad of analyses involved in producing our design.

Optimization of design elements was achieved using the built in numerical solver.

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