Design of a Light Business Jet Family

David C. Alman

AIAA : 498858

Andrew R. M. Hoeft

AIAA : 494351

Terry H. Ma

AIAA : 820228

Cameron B. McMillan

AIAA : 486025

Jagadeesh Movva

AIAA : 738175

Christopher L. Rolince

AIAA : 808866

I. Acknowledgements

We would like to thank Mr. Carl Johnson, Dr. Neil Weston, and the numerous Georgia Tech

faculty and students who have assisted in our personal and aerospace education, and this project

specifically. In addition, the authors would like to individually thank the following:

David C. Alman: My entire family, but in particular LCDR Allen E. Alman, USNR (BSAE

Purdue ’49) and father James D. Alman (BSAE Boston University ’87) for instilling in me a love

for aircraft, and Karrin B. Alman for being a wonderful mother and reading to me as a child. I’d

also like to thank my friends, including brother Mark T. Alman, who have provided advice, laughs,

and made life more fun. Also, I am forever indebted to Roe and Penny Stamps and the Stamps

President’s Scholarship Program for allowing me to attend Georgia Tech and to the Georgia Tech

Research Institute for providing me with incredible opportunities to learn and grow as an engineer.

Lastly, I’d like to thank the countless mentors who have believed in me, helped me learn, and

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provided the advice that has helped form who I am today.

Andrew R. M. Hoeft: As with every undertaking in my life, my involvement on this project

would not have been possible without the tireless support of my family and friends. I thank my

mother, Sue Ellen, for leading me through 23 years of life’s great challenges with enthusiasm, and

my stepfather, Martice, for helping me realize that I cannot know everything - and that’s okay. I

am especially grateful for my sister and brother-in-law, Maria and Cory: without them, I would

not have the privilege of attending Georgia Tech. And finally, I thank my longtime mentor and

friend, Dr. Kevin Eveker, for validating my interest in aerospace engineering as a profession and

providing uncountable hours of advice.

Terry H. Ma: I would like to thank my parents for teaching me early on the work ethics and

determination required to face and overcome the most daunting obstacles. Their endless support

for me has allowed me to attend Georgia Tech to further my education and passion for aviation and

engineering. In addition, I would like to thank the the faculty at Georgia Tech who has equipped me

with the skills needed to become successful, as well as the numerous talented students whom I have

had the pleasure of working along side. In particular, my close friend and colleague, Kayla W. has

been there for me personally and professionally throughout my 5 year career as an undergraduate

student at Georgia Tech.

Cameron B. McMillan My parents, Steve and Briggs McMillan, who have supported me on

every step of my educational journey. My brother, Andrew McMillan, for always grounding me.

I would especially like to thank Dr. Brian German and the many researchers of the Aerospace

Systems Design Lab at Georgia Tech for mentoring me throughout my undergraduate career.

Jagadeesh Movva My caring family, including my parents Ramesh and Kavitha Movva as

well as my brother Vikash Movva. Without their support to help me pursue my dreams, I would

not be where I am today. I am thankful for the numerous mentors I have had including Mr. Richard

Sims and Mr. Jason Weinberger. I would also like to thank my advisor, Prof. Marilyn Smith on

her guidance and mentorship.

Christopher L. Rolince: My incredibly supporting family, specifically my grandfather, Wayne

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L. Dolan, and father, Daniel J. Rolince for fostering a love and passion for aviation, and mother

Christine M. Rolince for supporting me throughout the rigors of Georgia Tech. Additionally,

thanks to the United States Navy for giving me the opportunity to both study at Georgia Tech and

continue my passion for flying. Additionally, Alyssa Bushhouse (Virginia Tech), Natalie Larkins

(Georgia Tech) and Alex Carroll (Georgia Tech) for their assistance with our team logo design and

model renderings.

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II. Executive Summary

The demand for light business jets fell significantly during and after the 2009 recession. Man-

ufacturers operating in this market have remained understandably cautious about introducing new

aircraft or technology. However, new investments in infrastructure coupled with global economic

growth in recent years present a high-return opportunity for new entries into the light business jet

market segment. This technical proposal outlines the design and development of a two-member

light business jet aircraft family with capacities of six and eight passengers. The goal of this project

was to introduce new technology and increased capability with a low operating cost in order to meet

current and future demands.

The eight-passenger variant was used as the primary design driver given its higher payload

and resulting performance impact. Multiple Class I concepts were considered; these included

conventional and unconventional component configurations across a variety of weight classes.

Ultimately, a canard-equipped aircraft with a maximum gross takeoff weight of roughly 17,000

lbs was selected. Rigorous technical analysis was performed to validate this design choice using

several commercial software packages for computational fluid dynamics, finite element analysis,

trade-study-guided optimization, and computer-aided design and simulation. In addition, several

custom software applications were developed in order to optimize the wing and canard size, loca-

tion, and stability characteristics. Due to the highly-coupled nature of the design, all analyses were

extensively iterated.

To guarantee the creation of a feasible light business jet family, both the six- and eight-passenger

variants were expected to share at least 70% parts commonality. Consequently, the six-passenger

aircraft demanded extensive design consideration to meet this requirement effectively. Ultimately,

a simple fuselage plug, six feet in length, was chosen as main differentiating element between the

two aircraft. While this choice complicated the center of gravity and stability characteristics to

achieve the desired performance, it significantly reduced the overall design complexity and created

a net reduction in development and acquisition costs. Since the wing and canard were sized to

meet the eight-person requirements, they are non-optimal for the six-person design mission. This

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does, however, allow the six-person to have a greater range and reduced cost due to increased parts

commonality and shared tooling. Performance and cost analyses validated this decision.

Financially, these clean-sheet aircraft represent a unique and lucrative opportunity for a new

or existing manufacturer in the light business jet market. Based on extensive cost modeling and

analysis, the current-year sale price of the eight-passenger aircraft was found to be US $7.5 million

while the current-year sale price of the six-person variant was found to be US $6.5 million. These

prices include a 15% profit margin for the manufacturer at a production rate of 6 aircraft per month.

The emphasis on feasibility and technology readiness throughout the design process results in high

confidence that these aircraft would meet a target entry-into-service date of 2020 and 2022 for the

six- and eight-passenger aircraft, respectively.

In light of the growing market potential and high financial upside, this light business jet aircraft

family shows substantial promise and constitutes a worthy business venture. Through the strategic

use of advanced yet proven technologies, commercial-off-the-shelf subsystems, and modern manu-

facturing practices, the HAMMMR Designs family of aircraft successfully delivers a sophisticated

aesthetic while accomplishing an impressive performance envelope and guaranteeing maximum

value for both customers and shareholders.

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Figure 1: Scaled Three-View Image of the HAMMMR Designs H-800

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Contents

I Acknowledgements i

II Executive Summary iv

III Requirements 1

IV Business Jet Market Analysis 2

V Class I Configuration Selections 4A Fuselage Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B Wing Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6C Landing Gear Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6D Engine Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7E Engine Location . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7F Engine Number . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8G Tail Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9H Configuration Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

VI Aircraft Weight and Sizing 11A Mission Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11B Constraint Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12C Weight Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14D Aircraft Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

VII Wing and Canard Design 20A Process Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20B Center of Gravity Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20C Wing Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24D Canard Design and Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31E Configuration Optimizer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

VIII Stability and Control System Design 40A Longitudinal Motion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41B Lateral Motion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43

IX Materials Selection and Structural Configuration 45A Fuselage Design and Structural Analysis . . . . . . . . . . . . . . . . . . . . . . . 46B Wing Structural Design and Analysis . . . . . . . . . . . . . . . . . . . . . . . . . 50C Design of Vertical Tail and Canard Structure . . . . . . . . . . . . . . . . . . . . . 54D Maneuvering Envelop and V-N Diagram . . . . . . . . . . . . . . . . . . . . . . . 55

X Landing Gear Placement and Design 58

XI Cabin Design 61

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XII Range 66

XIII Subsystem Selections 69A Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69B Avionics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70C Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72

XIV Alternative Variants 76A Cargo Variant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76B Electronic/Signals Intelligence . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76

XV Cost Estimation 77A Development and Production Cost . . . . . . . . . . . . . . . . . . . . . . . . . . 77B Operating Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79

XVI Conclusion 82

XVII Appendix 84

List of Figures1 Scaled Three-View Image of the HAMMMR Designs H-800 . . . . . . . . . . . vi2 FAA Data on Market Growth . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23 Standard T-Tail, Cruciform Tail, and V-Tail Configurations . . . . . . . . . . . . 94 Selected Component Configuration . . . . . . . . . . . . . . . . . . . . . . . . . 105 NBAA IFR Range Profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116 Energy based constraint sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . 137 Historical Aircraft Weight Regression (pictures not to scale) . . . . . . . . . . . 158 Trade Study of Cruise Conditions Effect on Gross Weight . . . . . . . . . . . . . 169 Trade Study of Cruise Conditions Effect on Design Mission Block Time . . . . . 1710 FLOPS Output Drag Polar at Mach 0.8, 43k ft . . . . . . . . . . . . . . . . . . . 1811 H-800 Payload Range Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 1912 H-600 Payload Range Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 1913 Location of Aircraft Datum . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2214 Component Weight Layout of 6 Passenger Variant . . . . . . . . . . . . . . . . . 2215 Component Weight Layout of 8 Passenger Variant . . . . . . . . . . . . . . . . . 2316 Center of Gravity Excursion Diagram for 8 Passenger Variant . . . . . . . . . . . 2417 Center of Gravity Excursion Diagram for 8 Passenger Variant . . . . . . . . . . . 2418 Section Lift Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2619 Section Drag Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2620 Surface Growth Rate of Wing Meshes . . . . . . . . . . . . . . . . . . . . . . . 2721 Boundary Layer Mesh Refinement . . . . . . . . . . . . . . . . . . . . . . . . . 2722 Relative Mach Number of Canard . . . . . . . . . . . . . . . . . . . . . . . . . 2823 Relative Mach Number of Main Wing . . . . . . . . . . . . . . . . . . . . . . . 2824 Planforms of Wing Surfaces. Controls are scaled correctly relative to surface, but

surfaces are not to scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

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25 Locations Considered for Wing and Canard Placement . . . . . . . . . . . . . . 3226 Canard Moment Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3327 Depiction of Aerodynamic Coupling with Canard . . . . . . . . . . . . . . . . . 3428 Aerodynamic Benefit of Canard and Wing Coupling . . . . . . . . . . . . . . . . 3429 Stall Strip for Canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3530 Configuration Optimizer Execution Process Flow . . . . . . . . . . . . . . . . . 3731 Potential 6 and 8 Person Static Margin Shifts with Maximum Parts Commonality 3832 Short-period Flying Qualities . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4233 Longitudinal SCAS Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . 4334 Block diagram of Lateral SCAS . . . . . . . . . . . . . . . . . . . . . . . . . . 4535 Overview of Aircraft Structural Layout . . . . . . . . . . . . . . . . . . . . . . . 4636 Structural Configuration and Material Selection for Aircraft . . . . . . . . . . . . 4737 Fuselage Construction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4838 Fuselage Panel Layup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4939 Simplified Fuselage Structural Model and Mesh . . . . . . . . . . . . . . . . . . 4940 IRF Plot for Pressurized Composite Fuselage . . . . . . . . . . . . . . . . . . . 4941 Wing-Fuselage Attachment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5042 Wing Spar to Fuselage Interface . . . . . . . . . . . . . . . . . . . . . . . . . . 5043 Wing Structural Model and Mesh . . . . . . . . . . . . . . . . . . . . . . . . . . 5144 Wing Internal Structure Mesh . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5245 Pressure Load Application on Bottom Surface of Wing . . . . . . . . . . . . . . 5346 Shape of Pressure Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . 5347 Wing Deformation under Load . . . . . . . . . . . . . . . . . . . . . . . . . . . 5448 Tail Structural Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5549 Tail Attachment to Empennage . . . . . . . . . . . . . . . . . . . . . . . . . . . 5550 Canard Articulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5551 V-N Diagram for 6 Passenger Variant . . . . . . . . . . . . . . . . . . . . . . . . 5652 V-N Diagram for 8 Passenger Variant . . . . . . . . . . . . . . . . . . . . . . . . 5753 Longitudinal Tip-over Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . 6054 Lateral Tip-over Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6055 Main Gear Retraction Scheme . . . . . . . . . . . . . . . . . . . . . . . . . . . 6056 Nose Gear Retraction Scheme . . . . . . . . . . . . . . . . . . . . . . . . . . . 6057 Operating Empty Weight versus Cabin Height for Selected FAR 23 Certified

Business Jets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6258 Maximum Range versus Cabin Length for Selected FAR 23 Certified Business Jets 6359 H-600 and H-800 Cabin Cross Section . . . . . . . . . . . . . . . . . . . . . . . 6360 Cabin Interior Rendering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6461 Cabin Interior Rendering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6462 H-600 Cabin Top View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6563 H-800 Cabin Top View . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6564 2,500 Nautical Mile Range Centered On New York . . . . . . . . . . . . . . . . 6665 2,500 Nautical Mile Range Centered On London . . . . . . . . . . . . . . . . . . 6766 Yearly Global Business Jet Flights . . . . . . . . . . . . . . . . . . . . . . . . . 6867 Williams FJ-44 Jet Turbine Engine [26] . . . . . . . . . . . . . . . . . . . . . . 6968 Fuel System Architecture . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

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69 Comparable Business Jet Flight Deck Layout . . . . . . . . . . . . . . . . . . . 7170 Weather Radar Display . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7271 Leading Edge Interior De-Icing Layout . . . . . . . . . . . . . . . . . . . . . . . 7372 Canard De-Icing System Using Electro-Expulsive Separation System . . . . . . . 7473 Business Jet Main Wing De-Icing System Layout . . . . . . . . . . . . . . . . . 7474 General Aircraft Lighting Layout . . . . . . . . . . . . . . . . . . . . . . . . . . 7575 Night Rendering of the HAMMMR H-800 Showing Lighting Arrangement . . . 7576 H-800 ELINT/SIGINT Variant (EC-800) with Embedded Antennas (no SAR tub) 7777 Revenue as Function of Aircraft Cost . . . . . . . . . . . . . . . . . . . . . . . . 7978 Acquisition Cost Breakdown for 360 Aircraft . . . . . . . . . . . . . . . . . . . 8079 Operating Cost Breakdown at 1000 Hours per Year . . . . . . . . . . . . . . . . 8180 Variation in Cost per Flight Four with Number of Hours Flown per Year . . . . . 8181 HAMMMR H-800 with Nighttime Operating Lights . . . . . . . . . . . . . . . . 84

List of TablesI Matrix of Requirement Compliance . . . . . . . . . . . . . . . . . . . . . . . 1II Fuselage Configuration Figure of Merit . . . . . . . . . . . . . . . . . . . . . 5III Wing Configuration Figure of Merit . . . . . . . . . . . . . . . . . . . . . . . 6IV Landing Gear Configuration Figure of Merit . . . . . . . . . . . . . . . . . . . 6V Engine Number Figure of Merit . . . . . . . . . . . . . . . . . . . . . . . . . 8VI Engine Failure Probability . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8VII Tail Configuration Figure of Merit . . . . . . . . . . . . . . . . . . . . . . . . 9VIII Mission Segments for NBAA IFR Profile. . . . . . . . . . . . . . . . . . . . . 12IX Constraint Sizing Outputs . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14X 8 Person Sizing Mission: Crew and Payload Weights . . . . . . . . . . . . . . 14XI 6 Person Sizing Mission: Crew and Payload Weights . . . . . . . . . . . . . . 15XII State of Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18XIII Component Weights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21XIV Comparison of Airfoil Series . . . . . . . . . . . . . . . . . . . . . . . . . . . 25XV Characteristics of Main Wing . . . . . . . . . . . . . . . . . . . . . . . . . . 30XVI Characteristics of Vertical Tail . . . . . . . . . . . . . . . . . . . . . . . . . . 30XVII Winglet Design Effectiveness [10] . . . . . . . . . . . . . . . . . . . . . . . . 31XVIII Characteristics of Canard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35XIX Configuration Optimizer Parameter Sweep . . . . . . . . . . . . . . . . . . . 36XX Configuration Optimization Priorities . . . . . . . . . . . . . . . . . . . . . . 39XXI Sensitivity Study on Level of Commonality Versus Optimal Configuration

Point for H-600 and H-800 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39XXII Final Configuration Placement . . . . . . . . . . . . . . . . . . . . . . . . . . 40XXIII Final Configuration Static Margins. . . . . . . . . . . . . . . . . . . . . . . . 40XXIV Longitudinal Flying Qualities . . . . . . . . . . . . . . . . . . . . . . . . . . 41XXV Natural Longitudinal Flying Qualities . . . . . . . . . . . . . . . . . . . . . . 42XXVI Lateral Flying Qualities . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44XXVII Natural Lateral Flying Qualities . . . . . . . . . . . . . . . . . . . . . . . . . 44XXVIII Smeared Properties of 24 ply Carbon Fiber Laminates [18] . . . . . . . . . . . 51

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XXIX Landing Gear Loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58XXX Landing Gear Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . 59XXXI Market Penetration Required for given Aircraft Production Level and Associ-

ated Unit Cost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78

NomenclatureList of AcronymsACP ANSYS Composite PrePostAIAA The American Institute of Aeronautics and AstronauticsAVL Athena Vortex LatticeCD Drag coefficientCFD Computational fluid dynamicsCFR Code of Federal RegulationsDAPCA Development and Production Costs for AircraftELINT Electronics intelligenceFAA Federal Aviation AdministrationFLOPS Flight Optimization SystemHF High FrequencyICAO International Civil Aviation OrganizationIRF Inverse Reserve FactorLED Light-emitting diodeMGTOW Maximum gross takeoff weightNACA National Advisory Committee for AeronauticsNASA National Aeronautics and Space AdministrationNBAA National Business Aviation AssociationRFP Request for proposalSAR Synthetic aperture radarSARPS Standards and Recommended PracticesSAS Stability augmentation systemSIGINT Signals intelligenceSST Menter’s Shear Stress Transport turbulence modelTCAS Traffic Collision Avoidance SystemTSFC Thrust-specific fuel consumptionUSD United States DollarVHF Very-High FrequencyVOR Very-High Frequency Omnidirectional RangeWAAS Wide Area Augmentation SystemList of Symbolsc̄ Mean Chord LengthΓw Wing Dihedral AngleΛc/4 Quarter-Chord Sweep Angleλw Wing Taper RatioAR Aspect Ratio

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cf Flap ChordCL,max Maximum Lift CoefficientCL Lift coefficientcr Rudder ChordCD0 Zero-Lift Drag CoefficientCG Center of Gravitye Oswald efficiency numberL/D Lift-to-drag ratioLE Leading EdgeS Main Wing Planform AreaSa Aileron Planform AreaSc Canard Planform AreaSf Flap Planform AreaSr Rudder Planform AreaVv Vertical Tail Volume RatioXcg Center of Gravity X-locationXNP Neutral Point X-locationList of Unitsfpm feet per minuteft feetin inchesKIAS knots indicated airspeednm nautical milespsf pounds per square footsq ft square feet

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III. Requirements

Table I: Matrix of Requirement Compliance

Description Requirements Compliance Section

- Six Eight Six Eight -

Max Cruise Speed Mach 0.85 (30,000 feet) Mach 0.85 Mach 0.85 VI.BRate of Climb 3,500 fpm (sea level) 3,500 fpm 3,500 fpm VI.B

Service Ceiling 45,000 ft 45,000 ft 45,000 ft VI.BTakeoff Length (MGTOW) 4,000 ft 4,000 ft 4,000 ft VI.B

Landing Length 3,600 ft 3,600 ft 3,600 ft VI.CMinimum Range (NBAA IFR) 2,500 nm + 100 nm alternate 2,890 nm 2,750 nm VI.D

Certification FAR 23 FAR 25 FAR 23 FAR 25 XIIIBaggage Capacity 500 pounds / 30 ft3 1,000 pounds / 60 ft3 500lbs / 60 ft3 1000lbs / 60 ft3 XIIPassenger Capacity 6 pax 8 pax 6 pax 8 pax XII

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IV. Business Jet Market Analysis

The demand for light business jets fell significantly during and after the 2009 recession. How-

ever, a resurgent economy and expected tax cuts and infrastructure investments, especially in the

United States, have triggered a renewed interest in business jets. Figure 2 shows FAA projections

for growth into the next decade. As indicated on the graph, the market is projected to grow by an

estimated 18% between 2017 and 2027. More specifically, data indicate that approximately 750

business jet aircraft will be purchased every year [1]. The HAMMMR series of light business jets

attempts to compete in capability and cost for roughly 300 of these deliveries.

Figure 2: FAA Data on Market Growth [2]

Aircraft in the light business jet market have and continue to offer a wide variety of capabilities,

features, aesthetics, benefits, and also drawbacks for a diverse customer base. Identifying the

salient characteristics desired by customers is a difficult endeavor. It varies widely based on the

customer’s unique situation and desired mission. There are numerous complex trades to be made

between capability, acquisition cost, manufacturability, operating cost, and many other factors. To

narrow the potential design space and clarify the relative importance assigned to each attribute, a

target customer with fixed preferences must be chosen.

The set of factors impacting target customer selection include general socioeconomic trends,

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potential market size and lucrativeness, and customer satisfaction with present market options

based on competitive analysis, among others. Ultimately, an aircraft that can maximize aesthetic

sophistication and offer best-in-class experience for the target customer while also delivering com-

petitive performance along critical dimensions - range, cruise speed, useful payload capacity, etc.

- will generate substantial market interest. An aircraft that can do this while guaranteeing lower

acquisition and operating costs, longer lifespan, and higher customer satisfaction can transform

a stagnant market and consistently outperform competitors. After extensive analysis of potential

customers, three markets were found to be of particular interest and value. These are: mid-range

chartered shuttle services; private, individual operators; and, corporate business executives.

While the mid-range chartered shuttle service providers would welcome a new entry into the

light business jet market to differentiate themselves from competitors offering similar services,

this market is relatively saturated with turboprop aircraft and is comparatively young. For a charter

service provider to be effective, their operating market must have sufficiently affluent customers,

especially high traffic congestion, and several cities in close proximity to enable several route

options. Unfortunately, few locales have the required elements to warrant a successful charter

service provider at present. However, this opportunity will continue to grow as population density

and commuting congestion in metro-areas increases. The design team determined this market to

be of secondary interest.

Similar rationale applies to the private, individual operator market. While many potential cus-

tomers may be sufficiently affluent to afford a private aircraft, those that can likely already have one

or have little interest in purchasing one. The market is relatively saturated with existing light busi-

ness jets and experiences strong competition from smaller, less-capable turboprop aircraft since the

expected range requirement is lower when the aircraft is flown exclusively for leisure. Again, this

market provides a tangential opportunity for infrequent, small volume sales, but was determined to

be of secondary interest.

By contrast, the corporate business executive market is thriving and presents several opportu-

nities for growth in the coming decades. In an increasingly globalized economy, many successful

Page 3

corporations have turned to jet ownership to offer their executives the flexibility that complements

their busy schedules, guarantee availability for critical business functions, and remain competitive

with other corporations who use their aircraft(s) for marketing. Business jets serve an important

role in both ensuring executive transportation availability and enticing new executives to join the

company during transition periods. Additionally, large corporations can afford the high acquisi-

tion prices of new aircraft and can generally be up-sold for custom interiors or paint schemes. This

represents a great opportunity for high volume, high margin sales of desirable aircraft. Therefore,

the design team decided to target the corporate business executive market.

With this market selection, the design team required an understanding of the target customers

expectations to deliver an aircraft that would be worthwhile. Business executives demand high

performance capabilities (particularly range and cruise speed), extensive cabin amenities, leading-

edge technology incorporated in the design and avionics, and significant curb appeal. The aircraft

must have several best-in-class features to impress potential customers and deliver long-term value

or high resale value to justify the decision to financially conscious shareholders.

Ultimately, the aircraft should be designed with both manufacturability and cabin comfort in

mind. In order to be competitive with similar 6- and 8-passenger business jets currently on the

market, the HAMMMR series of aircraft must have a cabin height and width of at least 5.5 feet. In

order to deliver a best-in-class experience, the design was given a cabin height and width of 5.75

feet.

Additionally, the design team desired an acquisition price near $8.5 million in order to remain

competitive with other aircraft in this class. Any increase in acquisition cost must be met with

a proportional decrease in operating cost or the inclusion of additional technologies that would

warrant paying a premium.

V. Class I Configuration Selections

The first steps in the design process involved comparing the importance of different character-

istics, or figures of merit, of the aircraft. The weighting factors, based on the stakeholder require-

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ments outlined previously, were devised to give importance to those features. A weighted Pugh

method was used to evaluate the different configuration choices. The results of this analysis are

shown in the selections below.

A. Fuselage Configuration

Table II: Fuselage Configuration Figure of Merit

Fuselage Weighting − − −

− − Circular Fuselage Elliptical Fuselage Area Ruled FuselageManufacturability 3.80 4.83 3.00 2.67Weight Savings 3.80 3.17 3.00 3.33

Aerodynamic Performance 3.80 3.17 3.50 5.00Passenger Comfort 4.20 3.50 4.00 2.67

Total Points − 57.13 52.90 53.00

For the fuselage, three design options were evaluated. The circular fuselage is fairly conven-

tional, and would be the standard baseline fuselage assumed on an aircraft. Its benefits include

ease of manufacturing and weight savings. An elliptical fuselage was also considered, as it would

provide an improvement in passenger comfort and profile drag at the expense of increased man-

ufacturing cost. Lastly, an area ruled fuselage was considered, as this design would provide the

most aerodynamic benefit of reducing the zero lift drag and compressibility drag. However, this

would be at the expense of passenger comfort and manufacturability. In order to mitigate cost and

increase comfort, a pure circular fuselage was chosen.

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B. Wing Configuration

Table III: Wing Configuration Figure of Merit

Wing Weighting − − −

− − Low Wing High Wing Mid WingStability 4.20 2.80 4.40 3.20

Landing Gear 3.20 4.00 2.40 3.20Maintenance 3.80 4.40 2.60 3.20

Passenger Comfort 3.20 3.40 3.40 2.80Aerodynamic Benefit 3.30 3.10 2.70 3.00

Structural Benefit 2.50 3.20 2.70 2.30

Total Points − 70.39 62.58 60.45

The wing configuration selection for this aircraft compares low wing, mid wing, and high wing

designs.

The high wing provides greater roll stability, however it would be difficult to design a landing

gear since the structure would not be able to fold into the aircraft. Additionally, the passenger

comfort inside would be compromised with the structure. Because of this, few aircraft in this class

use such a wing placement. The mid wing reconciles the landing gear problem to some degree.

However, it still imposes passenger comfort and structural complexity penalties.

Low wing configurations are typical in many of the business jets in this class size. There are

numerous benefits including landing gear length minimal interference with the passenger cabin.

Given these considerations, a low wing design was selected for this aircraft.

C. Landing Gear Configuration

Table IV: Landing Gear Configuration Figure of Merit

Landing Gear Weighting − − −

− − Tricycle Tandem QuadrupleComponent Complexity 4.80 4.60 2.80 2.00

Weight 3.80 4.00 3.60 2.20

Total Points − 32.68 24.32 15.96

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The landing gear for this aircraft is fairly conventional. A tricycle landing gear provides the

necessary taxi maneuverability for this class of aircraft. This design choice also has the greatest

weight savings and structural benefits. For these reasons, a tricycle gear was selected.

D. Engine Type

In this segment of aircraft, medium bypass turbofans are common. They provide lower cabin

noise and increased performance, especially speed, compared to turbo-prop aircraft. Addition-

ally, a turbo-jet cannot offer the same efficiency as a turbofan engine. Lastly, there are numerous

available engines that would provide the capability desired for this aircraft.

E. Engine Location

There were numerous options for the location of the engines. There were a number of trade-

offs between noise reduction and structural benefit. Since this is a luxury business jet, decreasing

passenger noise would be extremely beneficial during long flights.

Locating the engines on top of the wings would provide more engine noise, as there would be

less structure shielding the cabin. Placing the engines either under the wing or in the rear of the

aircraft would provide increased sound attenuation. However, a low-wing design would increase

the difficulty of a below-wing engine due to landing gear height and FOD concerns. As a result,

the engines are mounted on the rear of the aircraft.

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F. Engine Number

Table V: Engine Number Figure of Merit

Engine Number Weighting − −

− − 2 Engines 3 EnginesReliability 3.80 3.33 4.00

Maintenance Costs 3.40 4.33 2.17Fuel Efficiency 3.60 4.40 2.20Weight Benefits 3.40 4.33 2.17

Total Points − 57.97 37.85

Most business jets have two engines. The number of engines is related to probability of failure,

fuel burn, and maintenance requirements. Table VI shows the probability of failure for each engine

number choice.

Table VI: Engine Failure Probability

Airplane with: 1 Engine 2 Engines 3 Engines2 Engines 2Pef Pef

2 N/A3 Engines 3Pef 3Pef

2 Pef3

A single engine has an unacceptably low reliability, especially given the potential for over-

ocean flights. Since engines have certain parts that are replaced on a flight hour basis, having 3

engines would impose a higher operating cost. Additionally, the fuel consumption and weight of

the aircraft would increase. Thus, the two engine configuration was selected for this aircraft.

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G. Tail Configuration

Table VII: Tail Configuration Figure of Merit

Tail Weighting − − −

− − T-Tail V-Tail Cruciform CanardAerodynamic Benefit 4.20 3.40 2.80 2.80 4.00

Engine Placement Options 2.60 3.80 2.20 3.00 3.80

Total Points − 24.16 20.00 19.56 26.68

T-tail configurations are common in many of the business jets currently on the market. This

configuration allows for the engines to be placed on the empennage. Additionally, the higher T-

tail helps the control surfaces to avoid wingtip vortices. The effects are similar in a cruciform

configuration. However, both surfaces suffer compared to the other configurations in that they

have the largest wetted areas. This causes their drag values to be higher, as there is more induced

and parasitic drag in these configurations. The V-tail improves this configuration by reducing

the wetted area, however limits engine placement on the empennage. Figure 3 shows these three

configurations.

Figure 3: Standard T-Tail, Cruciform Tail, and V-Tail Configurations

The canard configuration provides even more utility, as it provides the least amount of trim

drag. The maximum trimmed lift coefficients for a canard aircraft are larger than that for a con-

ventional airplane. Additionally, the trimmed lift to drag ratios are higher with a canard aircraft.

These factors are of substantial benefit given that the mission of the aircraft is long-range cruise.

While the introduction of the canard introduces some aerodynamic and control complexity, these

Page 9

issues have been overcome in several aircraft such as the Beechcraft Starship and Piaggio P180.

The final selection for the H-800 and H-600 is the canard configuration.

H. Configuration Results

The configuration selection process yielded an aircraft with a circular fuselage, low wing, tri-

cycle landing gear, two rear-mounted engines, and a canard. The finalized configuration can be

seen in Figure 4.

Figure 4: Selected Component Configuration

Page 10

VI. Aircraft Weight and Sizing

A. Mission Analysis

After configuration selection, it was necessary to further refine the aircraft mission in order to

support preliminary weight analysis. It was determined that the National Business Aviation Asso-

ciation (NBAA) Instrument Flight Rules (IFR) profile with a 100 nm alternate, shown in Figure 5,

would provide a useful model given the type of aircraft being designed. This is the standard pro-

file used by the Business and Commercial Aviation Planning and Purchasing Handbook for the

standardization of performance quotes.

Figure 5: NBAA IFR Range Profile (not to scale)

The mission was divided up into 9 segments, with the parameters for each segment of the

mission listed in Table VIII. As later sections describe, a cruise altitude of 43,000 feet was chosen.

Range credits were also given for both the climb and descent phases of flight. A 2,500 nm range

would easily allow for travel from Los Angeles (LAX) to New York (JFK), a distance of 2,150 nm.

Page 11

Table VIII: Mission Segments for NBAA IFR Profile.

Name Distance (nm) Altitude (ft)

2500 nm Design Mission

Takeoff - -Climb 210 0 - 43,000Cruise 2180 43,000

Descent 110 43,000 - 0Land - -

100 nm Alternate Divert

Climb 14 0 - 18,000Cruise 44.2 18,000

Descent 41.8 18,000 - 0Land - -

B. Constraint Sizing

Based on the work of Mattingly, Heiser, and Daley [3], an energy based approach was used for

constraint analysis and mission sizing. This process is used to compare the required takeoff thrust

to weight ratio across a range of wing loadings. The first step was to break down the requirements

to define the most demanding mission segments that could potentially size our aircraft. These were

identified as:

• Max. speed in steady level flight

• Max. climb

• Max. sustained turn

• Service ceiling

• Takeoff

• Approach speed

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Using these six mission phases, the analysis was completed. This sizing process used the final

values of zero-lift drag (CD,0), aspect ratio, and span efficiency factor. The value of span efficiency

used was 0.84, which is similar for other modern business aircraft equipped with winglets. Ad-

ditionally, the approach speed was assumed to be 120 knots at a full payload, half fuel load. The

takeoff constraint is analyzed at MGTOW. The service ceiling and climb constraints are analyzed

at the speeds of best climb.

Figure 6: Energy based constraint sizing

Figure 6 shows that the service ceiling constraint and approach speed constraint are the limiting

factors for this aircraft. The approach speed constraint is one of the main drivers because of the

intentionally low CL,max assumed in this analysis. This was done so a simple flap system could

be used, which will reduce acquisition and maintenance costs. The service ceiling constraint was

directly from the RFP but a weight ratio equivalent to a half fuel load. Because the service ceiling

constraint was close to the cruise altitude, the cruise condition was also checked with a constraint

analysis. This showed that the airplane is able to meet the cruise condition even up to Mach .85

with the highest fuel load achievable after climbing to 43,000 ft. The other constraint near the

sizing point is the takeoff constraint. This is a difficult constraint to achieve because the 6,000 ft.

Page 13

density altitude assumed. This was chosen to ensure the aircraft can takeoff at near MGTOW at

high altitude airports, such as Aspen-Pitkin. Additionally, the H-600 was checked against these

constraints and found to meet all of them. Despite the decreased wing loading driving up the

required thrust to weight,the increased thrust to weight gained by using the same engine out weight

this.

The final results from the constraint are listed in Table IX. The thrust to weight and wing

loading are typical for other business jets in this size.

Table IX: Constraint Sizing Outputs

TS L/WTO WTO/S

0.37 71ps f

C. Weight Sizing

It was decided that in order to meet parts commonality and performance goals, the aircraft

would be sized for the eight-passenger variant (H-800) first. After this was complete, modifications

would be made to produce a six-person aircraft (H-600). The initial weight sizing for the H-800

was conducted using the Breguet Range and Endurance equations. This was solved for mass

fuel fraction based on notional speed, altitude, thrust-specific-fuel-consumption (TSFC) and lift-

to-drag (L/D) ratios. Other factors included in this analysis were the crew weight and payload

weight. These were determined based on market analysis and the requirements and their final

values are shown in Table X.

Table X: 8 Person Sizing Mission: Crew and Payload Weights

Type Crew Passengers Luggage Total

Quantity 1 4 4 −

Weight (lb) 225 205 100 −

Total Weight (lb) 200 820 400 1,445

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Table XI: 6 Person Sizing Mission: Crew and Payload Weights

Type Crew Passengers Luggage Total

Quantity 1 2 2 −

Weight (lb) 225 205 100 −

Total Weight (lb) 225 410 200 835

After these were defined, empty weight and max gross takeoff weight (MGTOW) could then

be solved for iteratively. The relation between these two numbers - Empty Weight and MGTOW,

was found using historical data on similar aircraft. The regression line used for this analysis can

be seen in Figure 7.

Figure 7: Historical Aircraft Weight Regression (pictures not to scale)

After initial sizing using Class 1 methods, NASA’s Flight Optimization System (FLOPS) was

used to increase the fidelity of the performance calculations. FLOPS was run to minimize the

gross takeoff weight of the H-800 aircraft that could perform its design mission described earlier.

A scaled down CF34-8C5 engine deck was used for the propulsion system model and was provided

Page 15

by our advisers. Additionally, the engine fuel flow was increased by 10% to account for the effects

of scaling between the CF34 and engines in the lower thrust class necessary for this aircraft. The

FLOPS tuning parameters were validated against a SyberJet SJ30 FLOPS model developed from

publicly available data[4]. Initial size estimates were used for the wing, fuselage, and control

surfaces and the design was iterated until the final configuration was achieved. Composites were

assumed to be used extensively throughout the aircraft to allow for significant weight savings. This

was especially important because it allowed the H-800 to be below the FAR 23 maximum weight.

More detail on the composite usage is described in a later section.

A number of trade studies were then conducted with variations in payload, cruise speed, range,

and cruise altitude in order to select an optimal design point given performance and business con-

siderations. An example trade study on altitude and cruise Mach number effect on MGTOW can

be seen in Figure 8. Figure 9 shows the effects of Mach number and altitude on the block travel

time. It was ultimately decided that a cruise altitude of 43,000 feet at a speed of Mach 0.8 would

provide the optimum balance of speed, comfort (partially a function of time), and weight.

Figure 8: Trade Study of Cruise Conditions Effect on Gross Weight

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Figure 9: Trade Study of Cruise Conditions Effect on Design Mission Block Time

Constructing the drag polar in FLOPS does require some assumptions about the aircraft. A

critical assumption during this process was that there would be no wave drag present. This would

require careful design in order to ensure all parts of the aircraft remain below their critical Mach

number. This requirement drove a large part of the computational fluid dynamics based aerody-

namics analysis. After numerous iterations, the FLOPS cruise drag polar took on the qualities

shown in Figure 10.

The final sizing outputs from FLOPS can be seen in Table XII. These values are very compet-

itive for aircraft in this class. Additionally, based on the thrust to weight ratio from the constraint

sizing process, the necessary engine size can be determined. This results in an engine with a 3,200

lb thrust capability. Due to the short timeline for this aircraft’s development before entry into

service, the Williams FJ44 was chosen as our engine. It is a flight proven engine, a variant has sig-

nificant additional thrust for any future growth, and used in other aircraft of this size. FLOPS also

outputs the landing field length required for the aircraft. Despite the low value of lift coefficient

Page 17

Figure 10: FLOPS Output Drag Polar at Mach 0.8, 43k ft

assumed, both the H-800 and H-600 met the desired landing field length of 3,600 ft at a typical

landing where most of the fuel has been used.

Table XII: State of Design

H-800 H-600

MGTOW Empty Weight CD,0 MGTOW Empty Weight CD,0

17,650 9,700 0.0202 15,750 8,800 0.0200

D. Aircraft Performance

Once the aircraft has been sized and the determination of initial performance characteristics

have been made, the aircraft performance was analyzed. Both the H-800 and H-600 payload range

diagrams were created to show how the aircraft range is affected by payload. The H-800 payload

range diagram is shown in Figure 11 and the H-600 diagram is shown in Figure 12.

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Figure 11: H-800 Payload Range Diagram

Figure 12: H-600 Payload Range Diagram

These diagrams were generated using FLOPS in fixed-vehicle mode. This assumed that no

extra fuel capacity was available for the H-800 beyond that needed for the design mission described

earlier. The H-600 uses the same fuel capacity as the H-800 but is MGTOW limited at higher

Page 19

payloads. As shown in Figure 12, the H-600 also meets the required 2,500 nm range with the

design payload listed in Table XI. Note, the payload range diagram payload does not include the

pilot weight. Although our aircraft was designed to cruise 2,500 nm at Mach 0.8 with the design

payload listed above, it does have more range when flown slower. For the fuel conscious user or

for general flights where time isn’t an issue, fuel operating cost savings can be achieved.

VII. Wing and Canard Design

A. Process Overview

The design of the wing and canard proved to be the most challenging part of the project. Mul-

tiple factors, such as takeoff rotation, stability, trim drag, and center of gravity excursion had to be

considered in order to design a safe, high performance aircraft. Since this part of the design was

so dependent on center of gravity analysis, it is included here in lieu of an independent section.

In order to optimize the design and placement of the lifting surfaces and center of gravity, many

coupled iterations were conducted using software suites such as AVL, FLOPS, and the team’s own

MATLAB and Python programs.

B. Center of Gravity Analysis

The location of the center of gravity and neutral point are extremely important. Static stability

margin is a function of the relative position of the two. Ensuring the aircraft had a low static

stability margin allowed for the aircraft to operate with minimal trim drag at cruise, translating

into a higher efficiency for the aircraft.

A major complication for the center of gravity analysis between the variants was the use of the

six-foot plug to differentiate between the H-600 and H-800. Additionally, the position of several

systems had a large effect on the overall design of the aircraft. The weight and balance calculations

for both aircraft were a large component the design tool discussed later in this report.

The component weights were obtained using a FLOPS based Class II design method for the

8 passenger variant. The 6 passenger variant was then analyzed in a fixed weight mode, which

Page 20

allowed part weights to be common between the two variants, except for the values that are reduced

between the two aircraft. As seen in Table XIII, the subsystem weights remained the same between

the two variants.

Table XIII: Component Weights

8-Passenger Location (ft) 6-Passenger Location (ft)Component Weight (lbs) X (ft) Y (ft) X (ft) Y (ft)

Air Conditioning 317 43.3 0.17 37.3 0.17Anti-Icing 75 38.3 1.7 32.3 0.17

Auxiliary Power 321 42.5 1.7 36.5 1.7Avionics 408 41 -0.5 35 -0.5Canard 141 5.3 -1 5.3 -1

Cargo Weight 250 41.7 -1.17 35.7 -1.17Crew and Baggage 225 6.25 0.33 6.25 0.33

Electrical 953 40 0 34.0 0Engine Oil 31 45.3 4.16 39.3 4.16

Engines 1019 45.3 4.16 39.3 4.16Fuel System 218 33.1 -1.5 27.1 -1.5Furnishings 1717 / 1287.75 30 0 22 0

Fuselage 1851 / 1150 28.8 0 25.8 0Fuselage Fuel 4370 / 4060 33.1 -2.5 27.1 -2.5

Hydraulics 227 38.3 -1.7 32.3 -1.7Instruments 181 3.75 -2.2 3.75 -2.2Main Gear 355.6 34.7 -2.2 28.7 -2.2Main Wing 1399 34.7 -2.2 28.7 -2.2Nose Gear 152.4 7.5 -2.17 7.5 -2.17

Passenger Baggage 120 / 160 41.7 -1.17 35.7 -1.17Passenger Service 125 33.3 -1 27.3 -1

Passengers 540 / 720 24.17 -1 18.17 -1Surface Controls 333 34.7 -1.08 28.7 -1.08Unusable Fuel 264 34.1 -2.5 28.1 -2.5Vertical Tail 164 41.7 6.25 35.7 6.25Wing Fuel 1612 34.7 -2.17 28.7 -2.17

Note that the datum referenced for the aircraft is shown in Figure 13, which is the same for

both variants.

Page 21

Figure 13: Location of Aircraft Datum

The center of gravity placement was conducted with reference to typical component weights

in other business jet aircraft [5]. This can be seen in both the 6 passenger an 8 passenger variants

below in Figures 14 and 15 respectively. Most of the weight difference between the variants is in

the cabin due to the plug removal or addition. This continues to allow the large parts commonality

between the two variant aircraft.

Class II moments of inertia were calculated next. These values were used in the stability

analysis of the aircraft, as well as the canard rotation constraint analysis that is described later in

this report.

Figure 14: Component Weight Layout of 6 Passenger Variant

Page 22

Figure 15: Component Weight Layout of 8 Passenger Variant

The fuel tank system and its placement placement, as seen in Figure 68 later in this report,

allowed for the center of gravity to be kept very close to the neutral point throughout the mission.

This led to a low static stability margin, which in turn leads to a decrease in the trim drag and

increase in efficiency of the aircraft. As mentioned in the design sweep tool, the design focused on

giving a larger weight for the 8 passenger variant compared to the 6 passenger variant. Thus the

center of gravity excursion was also smaller on the 8 passenger variant compared to the 6 passenger

variant. The center of gravity excursion diagrams are shown for both variants in Figures 16 and

17.

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Figure 16: Center of Gravity ExcursionDiagram for 8 Passenger Variant

Figure 17: Center of Gravity ExcursionDiagram for 8 Passenger Variant

C. Wing Design

The wing was designed for optimal performance within the outlined business jet mission. Con-

straint sizing required the wing loading to be 71ps f . A wing area was obtained using the aircraft

weight developed during weight sizing. The wing area was sized from the performance considera-

tions of the 8 person variant. The main wing and canard were kept in common for both aircraft, as

this reduces both the manufacturing and design cost. This will have a negative effect on the trim

drag on the 6 person, however this was deemed an acceptable trade-off due to the reduction in cost.

The aspect ratio desired for this aircraft was 8.5. In addition, from the Class II drag polar,

the lift-to-drag ratio was also determined. The next step was to select an airfoil that would could

provide the required lift-to-drag ratio. Table XIV compares the traits of the various airfoil series.

As seen in Table XIV, the use of a 6-series airfoil is typical for business jet applications. A

NACA 6-series airfoil was selected for this aircraft, as these are typical for operation in the high-

subsonic flight regime. These airfoils are used since they tend to promote laminar flow on larger

Page 24

Table XIV: Comparison of Airfoil Series [6]

Family Advantages Disadvantages Applications

4-SeriesGood Stall Characteristics Low Maximum Lift Coefficient General Aviation

Small Deviation of Center of Pressure Higher Drag Coefficient Horizontal TailsNot Sensitive to Surface Roughness Large Pitching Moment

5-Series

High Maximum Lift Coefficient Poor Stall Behaviour General AviationLow Pitching Moment Relatively High Drag Piston Aircraft

Not Sensitive to Surface Roughness CommutersBusiness Jets

6-Series

High Maximum Lift Coefficient High Non-Optimal Drag Piston FightersLow Optimal Drag High Pitching Moment Business Jets

Optimized for High Speed Poor Stall Behavior Jet TrainersSusceptible to Roughness Supersonic Jets

portions of the airfoil. This reduces the section drag of the airfoils compared to previous gen-

erations. Business jets spend most of their flying lives in a cruise configuration so any potential

penalties in off-design scenarios are of little concern. The surface roughness due to weather con-

ditions will also be addressed further in the subsystems portion of this report, as well as in the

canard section. The final airfoil selected was the NACA 66210. An additional motivation for this

particular airfoil being selected was that experimental data was available through the non-linear

transonic regime - a critical part of the flight envelope for this aircraft.

Page 25

Figure 18: Section Lift Coefficient [7] Figure 19: Section Drag Coefficient [7]

The lift-to-drag ratio was developed from the constraint sizing. As seen from Figure 18 the

necessary Cl value during cruise was achieved. In this analysis, all of the necessary lift generated

for the aircraft arose from the main wing. With future analysis and refinement, it would be possible

to reduce the size of the main wing as the canard provides additional lift on the aircraft.

The wing characteristics were determined through a combination of methods that included both

empirical and analytical methods. The thickness ratio and sweep angle were both determined using

empirical methods outlined by Dr. Roskam [5] for the initial selection. Computational analysis was

conducted with a 3-D wing model to verify that the critical mach number is above the high-speed

cruise speed the aircraft.

A critical mach number check was performed using Star-CCM+ CFD. This was done at the

high speed cruise speed, using standard atmosphere conditions. A mesh was developed to provide

accurate refinement in local regions. A K-ω SST mentor turbulence model was used. From litera-

Page 26

ture this model has shown good results for airfoils at low angles of attack [8]. This is an expected

flow regime for wings on light business jets, as they are typical smooth flow without excessive

angles of attack or reversed flow, especially in cruise. The mesh contained refinement on both

the leading edge and trailing edge. In addition, tip vortices were refined in the appropriate mesh

region. The resulting mesh is shown in Figure 20. As seen in Figure 21, the surface layer has been

modeled with a growth rate of 5-10% stretch rate as seen in literature [9].

Figure 20: Surface Growth Rate of Wing Meshes

Figure 21: Boundary Layer Mesh Refinement

Page 27

Figure 22: Relative Mach Numberof Canard

Figure 23: Relative Mach Number of MainWing

Based on these results, a thickness to chord ratio of 10% was selected, as well as a Λc/4 of

30° for the main wing and 35° for the canard. The selection of the canard sweep angle will be

discussed further in the canard portion of this report. A variable sweep angle aircraft was not con-

sidered for the HAMMMR business jet, as it would incur a large weight penalty for only marginal

performance improvement. In addition, the sweep angle also aligns with similar category aircraft.

The increased in overall wing weight due to swept design is considered in the FLOPS Class II

methods as a penalty.

The taper ratio, λW, selected was selected to generate the necessary area and span for the

aspect ratio defined in the generated Class II drag polar. This is within the typical ratios of other

business jets. Many of these aircraft maintain a straight taper ratio, as these save manufacturing

and development costs for the aircraft.

From the constraint sizing, the takeoff and landing CL requirements are 2.0 and 2.1 respec-

tively. The increment to the section lift coefficient needed for the harder constraint, the landing

requirement, was 1.16. Since the necessary ∆CL values for the flap system were low in our de-

sign, a plain flap system could be utilized on the aircraft. This was accomplished with a flap area

Page 28

as shown in Figure 24. The placement of the flaps defines the location of the aft spar, as this is

the flap attachment point. A flap deflection of 27° is found to be the necessary angle to meet the

CL requirement of 2.1. This was well within reasonable limits of the flap extension of other air-

craft. Utilizing plain flaps over a complex flap system was also of benefit in the estimated cost of

acquisition and operating.

The control surfaces on the main wing were determined through aileron area ratios of other

business jets. This led to the main wing planform shown in Figure 24. Note that the canard sizing

determination will be discussed in greater detail in the following section in this report.

The vertical tail was sized using volume ratios defined by similar type aircraft. This vertical

tail size has shown that it is sufficient in lateral stability. In addition, the control authority of the

rudder was determined using a similar volume-based empirical approach. A symmetrical NACA

0009 airfoil was selected for the vertical tail. The characteristics of the vertical tail are given in

Table XVI.

Figure 24: Planforms of Wing Surfaces. Controls are scaled correctly relative to surface, butsurfaces are not to scale

The wing dihedral, Γw, was determined through both ground clearance and stability consider-

ations. Dihedral provides increased lateral stability for both the spiral and Dutch roll modes. The

Page 29

degree for the dihedral angle, 2° was selected through comparison with other business jet aircraft

[5]. An incidence angle was determined through the values of the cruise drag, and was determined

to be 0 °. This was sufficient for the required lift and drag at cruise. Additionally, a lower inci-

dence angle is more comfortable for passengers as the floor is closer to level. Washout was used

on the wing to suppress tip stall. This is typical of most aircraft as it can be designed into the wing

without a significant increase in cost and offers performance improvements.

The final wing characteristics are listed in Table XVIII.

Table XV: Characteristics of Main Wing

Parameter Value Parameter ValueWing Area 260 ft2 Aspect Ratio 8.5

Airfoil NACA 66210 Λc/4 30°Root Chord 8.91 ft Γw 2°Half Span 22.5 ft Twist -0.1°

cf/c 0.3 Sa/S 0.105λW 0.35

Table XVI: Characteristics of Vertical Tail

Parameter Value Parameter ValueWing Area 81 ft2 Aspect Ratio 5.54

Airfoil NACA 0009 Λc/4 45°Root Chord 9.0 ft Vv 0.06Half Span 8.2 ft λW 0.38

cr/c 0.35 Sr/S 0.18

In addition, an analysis of the lifting efficiency was conducted. For this analysis, prior validated

high fidelity CFD results were compared to examine the benefits of various winglet types [10]. The

analysis done by Reddy was to compare a baseline Boeing 757 wing to various winglet configura-

tions. These included configurations with no winglet, a blended winglet, a scimitar winglet. The

results of this comparison are seen below in Table XVII below.

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Table XVII: Winglet Design Effectiveness [10]

Winglet Type ∆ CL (in %) ∆ CD (in %)Plain Wing 0% 0%

Blended Winglet 3.41% −4.43%Scimitar Winglet 6.23% −5.73%

With the low wing design that is selected for this aircraft, it was not possible to use a sharklet

style winglet while maintaining the 5° lateral clearance. The optimizing data that was generated

by Reddy was utilized for the selection of the winglet type. Leveraging data from wind tunnel ex-

periments run by NASA, the percentage improvement was better understood through the transonic

regime. From the results of the investigation, it was found that a blended winglet added to a wing

will lead to improvement in the order of 5% in the aerodynamic performance of the wing [11].

This was fed into our analysis by improving the span efficiency factor in FLOPS by 5%, leading

to an improvement in the overall efficiency of the aircraft with a slight reduction in weight.

D. Canard Design and Placement

A canard was used on this aircraft due to the increased performance it offers during cruise. This

allows for enhanced range and performance of the aircraft. It is, however, more complicated than a

more conventional aircraft. The canard design and placement optimization was determined using a

custom built tool that would create a design sweep in which stability margin was minimized. This

is explained further in an accompanying section in this report. The center of gravity of the aircraft

is tightly coupled in the canard placement, as the canard affects the neutral point of the aircraft. In

addition, the canard rotation was a constraint for the canard size.

The moment generated by the canard was a function of both the lift generated by the canard and

the moment arm due to its location. Thus, every canard location choice resulted in neutral point

analysis to ensure that the aircraft achieved the desired static and dynamic stability characteristics.

The wing placement was also a function of this neutral point and stability analysis. Figure 25

shows the areas considered for both wing and canard placement as well as the resulting neutral

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point locations. Structural constraints, outlined in more detail in Section IX precluded a wider

location selection.

Figure 25: Locations Considered for Wing and Canard Placement

A key advantage of canard aircraft is that they have lower trim drag than conventional aircraft

due to the canard contributing to lift. In order to see the benefits of this, however, a static stability

margin of near 0 was desired. This would allow the aircraft to fly with minimal control surface

deflection, resulting in lower drag in cruise. This would contribute to lower operating cost through

reduced fuel burn. The static stability margin can be calculated using Equation 1.

S taticMargin =xNP

c̄−

xcg

c̄(1)

A critical factor for the design of a canard equipped aircraft is its ability to rotate off the ground

during the takeoff roll. While conventional aircraft horizontal tail generates a downward moment

to push the rear of the vehicle down, a canard must produce an upwards moment to raise, or

pull, the nose of the aircraft up during takeoff. Figure 26 shows the factors that affect the canard

rotation. These include the lift and drag of both the wing and canard, engine Z-height, inertial

forces, and gravity. This relationship is tightly coupled with the center of gravity of the aircraft, as

it significantly alters the moment arm required for the canard to rotate the aircraft. Several of these

factors were swept in the design sweep tool that was created, as it allowed for the identification of

configurations that would be capable of pitching the aircraft nose during takeoff.

The canard control surface was designed to be fully articulating. This allows for the most

control authority for the canard during the takeoff rotation to allow for the aircraft to rotate. A

similar system is employed on the Saab JAS Grippen fighter aircraft [12]. This feature allows for

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Figure 26: Canard Moment Analysis

a sufficient CL value to be generated by the canard during the takeoff rotation. In addition, this

allows for the aircraft to trim to a lower drag configuration in flight compared to a fixed incidence

angle canard. As an added bonus, the articulating canard improves the maneuverability of the

aircraft, especially at low speeds where the control force contribution from dynamic pressure is

reduced.

Ultimately, the aircraft’s canard needed to generate an upwards pitching moment of sufficient

magnitude to allow for a rotation rate of 3-4 degrees/second [13]. This rotation rate is seen to be

typical in aircraft design, as it ensures that the tail of the aircraft clears the runway. In addition, it is

an accepted design value for the comfort of passengers on-board. The rotation rate was calculated

using the moment of inertia along the location of the main gear, Izz*. The moment of inertia about

the center of gravity was calculated using the Class II method mentioned in the prior portion of this

report, this the Izz is converted to the moment of inertia about the main gear contact point through

parallel axis theorem, resulting in Izz*. Utilizing this moment of inertia, it was possible to calculate

the rotation rate based on the forces that are at play on the aircraft.

Another benefit of the canard configuration is the resultant improvement of the max lift coef-

ficient of the main wing. Thus the canard was placed higher on the fuselage as compared to the

main wing so the trailing vortices would improve the flowfield on the upper surface of the main

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wing. An image of this effect on the Saab JAS Grippen is shown in Figure 27. The resultant effect

on the lift coefficients is seen in Figure 28. In the analysis performed, this improvement could

not be modeled with the computational resources at hand. Thus, this improvement factor was not

incorporated in the performed analysis.

Figure 27: Depiction of AerodynamicCoupling with Canard [5]

Figure 28: Aerodynamic Benefit of Ca-nard and Wing Coupling [5]

A major consideration for canard aircraft is the prevention of deep stall. This would likely be

catastrophic. A deep stall would be avoidable if the canard stalled first on the aircraft, causing the

nose to drop and angle-of-attack to decrease. The design of the canard has an increase in sweep that

would ensure it stalled before the wing. The stability and control augmentation system would pre-

vent the aircraft from entering a deep stall condition as well. This would be accomplished through

control input limitation and stall detection. In addition, leading edge devices were implemented

on a short inboard portion of the canard to hasten the stall of the canard. The design of the leading

edge device is shown in Figure 29.

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Figure 29: Stall Strip for Canard [14]

Through the design methods employed on the constraint of the canard, the final canard design

specifications are shown in Table XVIII

Table XVIII: Characteristics of Canard

Parameter Value Parameter ValueWing Area 54.7 ft2 Aspect Ratio 5.62

Airfoil NACA 66210 Λc/4 35°Root Chord 4 ft Γw 0°

Span 8.4 ft Twist 0°λW 0.5 Se/S 0.87

E. Configuration Optimizer

The team developed a tool in Python to test multiple component location and size variations and

their effect on the static margins, CGs, and rotation rates for the two aircraft variants. This tool ran

a linear search of six design parameters and output the static margins, CGs, and rotation rates for

the two variants. These were analyzed for the worst case condition for both aircraft. Wet and dry

CGs were analyzed so the in-flight static margin shift could be predicted. JMP, a statistics-heavy

mathematical software package, was then used to create surrogate models of the output variables.

These models were then used in an optimization to minimize trimmed drag.

The first step in the tool development process was to determine which design parameters should

be varied and what range of values they should be tested across. It was determined that the wing

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location, engine location, fuselage fuel location, canard area, and canard location should be varied

since these parameters had the most effect on the output variables being predicted. Varying the

vertical tail location was considered but excluded because the tail would have needed to be resized

in every loop, driving up code complexity and run time. The parameters varied and the range

of values considered are shown in Table XIX. A full factorial experiment was conducted for the

variables in this design space. To reduce the computation time of the evaluation process, parallel

processing was utilized. The frame of reference used for these measurements is the origin is at the

fuselage centerline at the same longitudinal location as the nose of each aircraft. The x-axis points

back along the fuselage centerline, and the z-axis is upward towards the roof of the cabin.

Table XIX: Configuration Optimizer Parameter Sweep

Parameter Min. Value Max. Value Number Tested

Wing Root LE (ft) 25.0 38.2 9Engine X-location (ft) 38.17 45.83 9Engine Z-location (ft) 1.25 5.0 9

Fuselage Fuel x-location (ft.) 31.25 43.33 9Canard x-location (ft) 0.0 7.0 9

Canard Area (ft2) 40 60 9

Once the design space was identified, the overall tool execution process was laid out. This

is shown in Figure 30. This process has two parallel paths. First, Athena Vortex Lattice (AVL)

[15]was used to calculate the neutral point of the aircraft. Initially, this was inside of the code

loop, but the individual case run-time was too slow for reasonable analysis of the whole trade

space. A surrogate model was generated for calculating the neutral point in order to speed up

execution time. This model was created using JMP neural net modeling based on the design space

parameters listed in Table XIX for wing location, canard location, and canard area. Canard aspect

ratio, taper ratio, and sweep angle were held constant as the area varied. This was chosen so

the changing canard area had less impact on the performance calculations. Once this model was

generated, it was used in the configuration optimizer and significantly sped up the execution time.

The other parallel branch was to calculate the center-of-gravity and then calculate the rotation

Page 36

rate for the canard. These two processes leverage much of the analysis done in other sections on

canard design and CG analysis.

Figure 30: Configuration Optimizer Execution Process Flow

Once the code flow was designed, the tool was created and results were generated. The impor-

tant check was to see whether static margin shifts of zero could be attained for both variants with

no major design changes between the two variants except for the plug. The results of this analysis

are shown in Figure 31. It is possible to have static margin shifts of less than 5% for both variants

with no changes between the two except the fuselage plug.

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Figure 31: Potential 6 and 8 Person Static Margin Shifts with Maximum Parts Commonality

Using JMP neural net modeling, surrogate models were created for the static margins of the

two aircraft variants. An optimization was then performed using JMP to minimize trimmed drag

and center-of-gravity shift between the fueled and unfueled condition. The optimization priorities

are shown in Table XX. The H-800 variant was given a 2:1 priority over the H-600 and the same

desirability distribution was set around each parameter. This weighting was the result of market

analysis that expected the H-800 to be sold in larger numbers and generate more revenue than the

H-600.

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Table XX: Configuration Optimization Priorities

Parameter Goal Value

H-600

6 Person Wet Static Margin 0%6 Person Dry Static Margin 0%6 Person Static Margin Shift 0%

H-800

8 Person Wet Static Margin 0%8 Person Dry Static Margin 0%8 Person Static Margin Shift 0%

Once the optimization setup was created, three optimizations were run to study the sensitivities

of the parts commonality on the final design point. The three scenarios considered were only

same canard, wing, and fuel locations for both variants, changing only the fuel center-of-gravity

location between the two, and changing the fuel center-of-gravity and the canard area between the

two variants. The results of this are shown in Table XXI.

Table XXI: Sensitivity Study on Level of Commonality Versus Optimal Configuration Point forH-600 and H-800

Variant Wet Static Margin (%) Dry Static Margin (%) Static Margin Shift (%)

Only Plug

H-600 6.6 3.3 3.2H-800 −3.2 −1.3 −1.9

Plug + Fuselage Fuel Change

H-600 0.9 2.9 −2.0H-800 0.1 0.2 0.1

Plug + Fuselage Fuel + Canard Area

H-600 0.0 0.0 0.0H-800 0.0 0.0 0.0

As shown in Table XXI, increasing the level of parts difference between the two variants de-

creases the static margins and static margin shifts. Based on these results, a design with only the

plug and fuselage fuel location changes was selected. This was chosen because the fuselage fuel

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center-of-gravity location is easily achievable by pumping fuel between our existing fuel tanks and

gives a good trimmed drag benefit. The canard area change was not included because the small

benefit it added was not worth the decrease in parts commonality.

Tables XXII and XXIII show the final design point and static margins for both variants. Due to

the plug length, the H-600 x-locations behind where the plug would be are shifted six feet forwards

relative to the H-800. This table shows that both variants have very low static margins and static

margin shifts. This was our design goal from the beginning with the canard and was achieved. The

small static margin shift for the the H-600 was deemed acceptable since the aircraft was optimized

for the H-800 and some lowered performance was expected.

Table XXII: Final Configuration Placement

Canard Area (ft2) Canard X (ft) Wing X (ft) Engine X (ft) Engine Y (ft) Fuselage Fuel X (ft)

H-600 54.7 0 25 45.8 4.2 33.1H-800 54.7 0 19 39.8 4.2 28.9

Table XXIII: Final Configuration Static Margins.

Variant Wet Static Margin (%) Dry Static Margin (%) Static Margin Shift (%)

H-600 0.9 2.9 −2.0H-800 0.1 0.2 0.1

VIII. Stability and Control System Design

Before finalizing the design point, it was also necessary to analyze the dynamic stability of the

aircraft. It was determined that a preliminary stability and control augmentation system (SCAS)

design would confirm that the aircraft is controllable. This control system design followed the

paradigm of classic controls with fixed gain values. A modern controls approach would allow for

even better handling qualities. Analysis was conducted for both longitudinal and lateral motion.

Since passenger comfort was a driving design objective, it as important to not only have a

stable aircraft, but one that was comfortable. While the relationship between passenger comfort

and handling qualities is outside the scope of this project, a full control system design would

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include heavy input from test subjects to ensure comfortability.

A. Longitudinal Motion

It was desired that the aircraft attain Level I handling qualities - defined on the Cooper-Harper

scale as, "excellent, highly desirable," with pilot compensation not a factor for desired perfor-

mance. The requirements to achieve this for long and short-period longitudinal motion are shown

in Table XXIV. The short-period mode requirements can also be seen visually in Figure 32.

Table XXIV: Longitudinal Flying Qualities

Phugoid ModeLevel 1 ζ > 0.04Level 2 ζ > 0Level 3 T2 > 55s

Short-period Modeζsp ωn, sp

min. max. min. max.Level 1 0.4 1.3 2.4 3.8Level 2 0.25 1.3 1.8 6.3Level 3 0.15 − 1.8 −

In order to analyze these modes, it was necessary to populate a state space model of the aircraft

with the appropriate stability derivatives. AVL provided stability derivatives as an output of its

analysis, so these were easily inputted into the model. Equations 2 - 5 show a representative

buildup of the state space model for longitudinal stability analysis [16].

Alon =

Xum

Xwm 0 −gcosθ0

Zum−Zω̇

Zwm−Zω̇

Zq+mU0

m−Zω̇−mgsinθ0

m−Zω̇

I−1yy [Mu + ZuΓ] I−1

yy [Mw + ZwΓ] I−1yy [Mq + (Zq + mU0)Γ] −I−1

yy mgsinθ0Γ

0 0 1 0

(2)

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Figure 32: Short-period Flying Qualities [16]

Blon =

Xδem

Xδp

m

Zδem−Zω̇

Zδp

m−Zω̇

I−1yy [Mδe + ZδeΓ] I−1

yy [Mδe + ZδeΓ]

0 0

(3)

x′ = Alon xlon + Blonulon (4)

y = Clon xlon + Dlonulon (5)

Table XXV shows the results of the natural eigenvalue analysis.

Table XXV: Natural Longitudinal Flying Qualities

Longitudinal Flying QualitiesMode Level 1 Requirement Result Analysis

Phugoid ζ > 0.04 ζ = 0.0087 Unsatisfactory

Short-Period0.4 < ζ < 1.3 ζ = 0.36 Unsatisfactory2.4 < ω < 3.8 ω = 2.99 Satisfactory

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This analysis showed that neither the Phugoid or Short Period modes fully met Level I handling

qualities, even though they were stable. This was common across both variants. To resolve this

issue, a root-locus plot was analyzed and manipulated to achieve the desired characteristics. Per

most longitudinal controller designs, both angle of attack and pitch rate feedback were used for

the controller. This moved the damping ratios and natural frequencies into Level I ranges. Only

small gains on the controller were necessary since the system was already close to meeting the

requirements without the addition of a controller. A block diagram of the model, as developed in

Simulink, can be seen in Figure 33.

Figure 33: Longitudinal SCAS Block Diagram

B. Lateral Motion

Level I handling qualities were also desired for lateral motion. The requirements to achieve

this for spiral mode, roll mode, and dutch roll can be seen in Table XXVI.

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Table XXVI: Lateral Flying Qualities

Roll Mode Time ConstantLevel 1 20sLevel 2 12sLevel 3 4s

Spiral Mode Minimum Doubling TimeLevel 1 12sLevel 2 8sLevel 3 4sDutch Roll Modeζmin ωn,min ζ ∗ ωn

Level 1 0.19 1 0.35Level 2 0.02 0.4 0.05Level 3 0.02 0.04 No Limit

Once again, it was necessary to populate a state space model of the aircraft with the appropri-

ate stability derivatives as provided by AVL. Eigenvalue analysis was again conducted to see the

aircraft’s natural lateral flying qualities. The results of this are shown in Table XXVII.

Table XXVII: Natural Lateral Flying Qualities

Lateral Flying QualitiesMode Level 1 Requirement Result AnalysisRoll t < 20s 7.38s Unsatisfactory

Spiral t < 12s 12.63s Satisfactory

Dutch Rollζ > 0.19 0.0038 Unsatisfactoryωn > 1 0.23 Unsatisfactory

ζ ∗ ωn > 0.35 0.001 Unsatisfactory

As the table shows, both roll and dutch-roll modes are stable but do not meet the Level I

handling qualities. This was common across both variants, with the H-600 exhibiting slightly

worse characteristics. In order to compensate for this, a lateral-stability controller was designed

with yaw and roll-rate feedback. This allowed the aircraft to meet Level I handling requirements -

essential for passenger comfort. One issue that arose was that the yaw rate feedback is non-zero in

a steady-state turn. In order to eliminate the unwanted steady state feedback of yaw rate to rudder

deflection during a turn, it was also necessary to implement a high-pass (washout) filter.

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A block diagram of the controller, as developed in Simulink, can be seen in Figure 34.

Figure 34: Block diagram of Lateral SCAS

IX. Materials Selection and Structural Configuration

Early in the design process, a primary design goal was to increase aircraft performance and

parts commonality while keeping the same or lowering aircraft weight and thus operating costs.

As a result, the HAMMMR series incorporates composite materials in many of the structural ele-

ments. The composite materials used in this aircraft are primarily carbon fiber/epoxy resin systems

and sandwich structures. The strands of carbon fiber embedded in an epoxy resin matrix are formed

into thin sheets known as laminates. When assembled into layers, they can have very high strength

to weight characteristics. In addition the lay up, or the stacking angle sequence between the lay-

ers, can be adjusted in order to strengthen the structure in a particular direction. This is due to

the orthotropic nature of the carbon fiber material - exhibiting different properties (strength and

stiffness) along each orthogonal axes of the material. This one of the advantages of composite

materials over conventional metallic structures (such that those utilizing aluminum alloys or Tita-

nium). Composite structures can be designed and optimized to meet both high strength and lower

Page 45

Figure 35: Overview of Aircraft Structural Layout

weight requirements. This section outlines material utilization choices, weight reduction estimates,

and the structural analyses of selected composites.

A material breakdown by primary structural components and estimates of weight saved over a

conventional metallic construction can be seen in Figure 36. While this construction is considered

representative of the aircraft design, optimization on ply thicknesses, ply shapes, as well as specific

laminate properties can be completed to further reduce aircraft weight and maximize the material

usage efficiency. The structural analyses was primarily conducted using ANSYS Mechanical (fi-

nite element solver).

A. Fuselage Design and Structural Analysis

The fuselage of the aircraft is constructed using a combination of composite sandwich panels

and carbon fiber based stiffeners, doubling as flanges and interior structure. This design drastically

reduces the amount of formers (stiffening elements) required on the fuselage. The panels are

partitioned into quarters at the 12, 3, 6 and 9 o’clock positions circumferentially and are joined

together using Titanium bolts or blind rivets. The 6-passenger and 8-passenger variants of the

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Figure 36: Structural Configuration and Material Selection for Aircraft (Target Weight ReductionValues Recommended by Roskam [17])

aircraft have 3 and 4 longitudinal sections, respectively. This modular design means that any

panels damaged during operation or maintenance work can easily be removed and replaced with

spares. This reduces aircraft downtime due to fuselage repairs - improving revenue streams - and

assists the operator in spare parts management and stock disposition.

The panels are constructed using a honeycomb core with unidirectional carbon fiber/epoxy

matrix prepreg laminates. The stiffness capabilities of the panels reduced the number of formers

and fully eliminated longerons used in traditional metallic designs to further reduce weight. The

honeycomb is comprised of hexagonal cells and provides high out-of-plane stiffness. For in-plane

stiffness as well as damage tolerance capability, face sheets are bonded to the honeycomb core and

co-cured to form each panel. Additionally, a metallic mesh was incorporated into the laminate for

electrical bonding and lightning protection. The honeycomb core thickness is 0.125 inch and was

bonded to 8 plies of face sheet laid up quasi-isotropically in [0/45/-45/90] degrees on both sides

of the core. Each face sheet is 0.0085 inch thick. Figures 37 and 38 details the internal fuselage

structure and panel layup. The stiffeners are constructed using layers of unidirectional carbon

Page 47

fiber/epoxy matrix laminates.

Figure 37: Fuselage Construction

The structure of the fuselage was analyzed for cabin pressurization to sea level at design cruise

altitude of 43,000 feet. A pressure load accounting for the exterior and interior pressure difference

was applied to the inner wall of the fuselage. For structural analysis modeling purposes, the ge-

ometry of the fuselage was reduced to the cabin section with the formers/flanges modeled as rings

with the same width and thickness as the actual design. The geometry and structural mesh is shown

in Figure 39. Additionally, the ANSYS Composite PrePost (ACP) is used in order to model the

composite plies and core. In this analysis, generic orthotropic material properties provided by the

ANSYS materials library were used. The honeycomb was modeled as a low density material, with

out of plane stiffness much greater than the in plane stiffness. Each facesheet is unidirectional with

higher stiffness in the fiber direction than the transverse direction. The panel and formers were

meshed and modeled as extruded SHELL181 elements, with bonded ("glued") contacts between

each panel and from the panel to its corresponding flange. The results of the pressurization analysis

is shown in Figure 40. Using the Tsai-Wu failure criterion built into the ANSYS materials library,

the Inverse Reserve Factor (IRF) is shown with considerable margin to failure as the IRF is much

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less than 1 (where IRF = 1 indicates failure).

Figure 38: Fuselage Panel Layup

Figure 39: Simplified Fuselage Structural Model and Mesh

Figure 40: IRF Plot for Pressurized Composite Fuselage

Page 49

B. Wing Structural Design and Analysis

The wing is a cantilever design in order to eliminate additional drag which would be introduced

with external attachment structures such as braces or struts. The swept wing was connected to the

fuselage via the dual spars with I-beam cross section. The I-beam provides stiffness against bend-

ing as well as shear loads from aerodynamic and inertial forces. They protrude outwards from the

root, and are bent to accommodate the sweep angle. The I-beams were reinforced with metallic

end plates fastened onto a box beam spanning between two formers. Figure 42 detail the interface

between the spar and the fuselage structure. The box beam was fastened to the formers at each end

and provide warping and torsional load transfer from the spars to the fuselage. On the wing lie two

primary control surfaces - ailerons and plain flaps. Both the ailerons and flaps are actuated using

hydraulic motors embedded in the internal structure of the wing, and geared to their respective

shafts. With a pilot input of control yoke/stick or flap deployment, the pressure in the hydraulic

lines is regulated and control surfaces are deflected to the desired positions.

Figure 41: Wing-Fuselage Attachment Figure 42: Wing Spar to Fuselage Interface

The wing is made of a carbon fiber and foam core skin with dual spars and ribs constructed

with carbon fiber/epoxy matrix. The skin provides an aerodynamic surface, while the ribs hold up

the airfoil shape of the skin. The spars are the primary load carrying members. The skin is made

Page 50

up of a 0.4 inch thick foam core, and 3 plies of 0.0085 inch thick plain weave carbon fiber prepreg

oriented in [0/60/0] degrees. The ribs are uniform in thickness, and are 2 inches thick, with 12

ribs spaced evenly across the span. The front and rear spars are placed at 15% and 75% chord,

respectively, with the rear spar coincident with the aileron and flap positions. The structure is then

cured as a whole. The rib and spar material properties used in the analysis were representative of

a 24-ply quasi-isotropic carbon fiber layup with smeared properties given in Table XXVIII. The

specific ply layups were not modeled due to the complexity in the spar cross-sectional geometry.

The skin was modeled with ANSYS material properties as described.

Table XXVIII: Smeared Properties of 24 ply Carbon Fiber Laminates [18]

Young’s Modulus 7861.045 ksiPoisson’s Ratio 0.305Shear Modulus 3016.785 ksi

The wing skin was modeled using SHELL181 elements (derived from ply/core model); the

ribs and spars were modeled as SOLID186 elements. To model the co-cured components, bonded

contacts were applied between the upper and lower skin surface and between the skin, ribs, and

spars. The wing was then constrained by applying a fixed support on the end plate of the spar at

the root. Loads were applied in the form of gravity and pressure distribution.

Figure 43: Wing Structural Model and Mesh

Page 51

Figure 44: Wing Internal Structure Mesh

Due to the concurrent design effort with the flight dynamics model and 3-dimensional unsteady

computational fluid dynamics, accurate pressure loads were not available for the analysis. Instead,

an elliptical lift distribution was calculated along the leading edge of the wing and uniform in the

chord, and is shown in Figures 45 and 46. The equation for the lift distribution [19] is:

q(z) =2WTOn

√L2 − z2

piL2 ∗ 2 ∗ [(Cr −Ct)/b + Cr](6)

Applying the pressure load distribution normal to the bottom of the wing surface resulted in

a pressure distribution uniform along the chord, and approximately elliptical along the span. The

pressure load was obtained by using a load factor of 3, which is the upper limit of the maneuvering

load envelope. Because of the simplifying assumptions made in the materials and its modeling,

failure criterion was not computed. Instead, deflection results are shown in order to visualize and

baseline the elasticity of the wing. Note that the result in Figure 47 was amplified by 1.5 times

in order to differentiate the deformed and undeformed geometry. A maximum deflection of 5.1

inches is computed at the tip of the wing.

Page 52

Figure 45: Pressure Load Application on Bottom Surface of Wing

Figure 46: Shape of Pressure Distribution

Page 53

Figure 47: Wing Deformation under Load Factor n = 3

C. Design of Vertical Tail and Canard Structure

The remaining lifting surfaces, canard and vertical tail, share a very similar internal structure to

the main wing. The vertical tail and rudder were sized as described previously, and were comprised

of skin, ribs, and I cross section spars. The ribs and the spars were, as in the wing, developed using

carbon fiber materials. The skin is a foam core skin with carbon fiber prepreg facesheets for the

additional stiffness. As the tail undergoes less severe loading than the main wings, the stiffness

of the tail could be reduced in order to achieve additional weight savings. The spars are placed at

15% and 65% chord, with the rear spar coincident with the rudder position. The end of the spars

are bent in order to accommodate the back swept design. They are then connected to empennage

structural elements via bolts. Much like the aileron, the rudder is also controlled via hydraulic

power. Pedal position instead is used to regulate hydraulic pressure in order to turn a hydraulic

motor. The motor is placed inside the tail and geared to the shaft along which the rudder lies.

The canard is, structurally, a scaled model of the main wing, with the aerodynamic, geometric,

and cross sectional properties defined previously. The materials used were the same as those used

for the wing. The canard was split into two sections - the inboard section which is fixed to the

fuselage, and the outboard section which can freely pivot and be controlled via hydraulic motor

actuators. Each canard is designed for independent articulation. This allows for greater control

Page 54

Figure 48: Tail Structural Layout Figure 49: Tail Attachment to Empen-nage

authority and flexibility.

Figure 50: Canard Articulation

D. Maneuvering Envelop and V-N Diagram

In determining the load factors and gust loads, 14 CFR 23.337 (normal category) and 14 CFR

25.337 were consulted for both the H-600 and H-800. Following the process outlined in Roskam

Part V [20], the stall, maneuvering, design cruise, and design dive speeds are calculated or esti-

Page 55

mated. By establishing a positive limit load factor of 3 and negative limit load factor of -1, the V-N

diagrams with gust loads are presented in Figures 51 and 52.

Figure 51: V-N Diagram for 6 Passenger Variant

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Figure 52: V-N Diagram for 8 Passenger Variant

Page 57

X. Landing Gear Placement and Design

The landing gear is often a critical part of the design process. The methodology of this design

was primarily based on a desire to meet longitudinal and lateral tip-over and ground clearance

criterion. It was hoped to meet this while maintaining low operating cost through the use of a

low-complexity design. Additionally, the landing gear placement was an iterative process - tightly

coupled with the center of gravity analysis - which was in turn highly coupled with the canard and

wing placement. To ease the computational burden, the landing gear placement was left outside

the loop described above. Instead, empirical data was used to narrow the possible landing gear

placement options. After the center of gravity and final weight had been determined, the landing

gear was checked to ensure it met requirements. The first part of the process was to calculate the

load on each strut. In order to do this, the location of the gear had to be determined. Figure XXX

shows the location of the landing gear. Table XXIX shows the results of the load calculations.

Table XXIX: Landing Gear Loads

- H-800 H-600

- Weight (lbs) Percent Weight (lbs) Percent

Nose Gear 1,091 6.2% 813 5.3%Main Gear 16,549 93.9% 14,417 94.7%

It was first necessary to check that the weight percentage on the nose gear and main gear was

appropriate. Typically, business jets have between 5% and 10% of the aircraft’s weight on the

nose gear. Calculations showed that this was the case for the H-800 as well. The H-600, with

the plug removed, had slightly more weight on the main gear but was still within margins. After

determining the load on each gear, the number of tires could be selected. For both aircraft, it was

concluded that two tires were needed per main strut. Only one tire was necessary for the front

gear. These determinations were based off of data provided in Roskam (insert citation here). Some

useful metrics for each landing gear can be seen in Table XXX.

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Table XXX: Landing Gear Characteristics

- Front Gear Main Gear

Number of Tires 1 2Tire Size (in) 18x6 25x8

Tire Pressure (psi) 110 135

A critical aspect of the landing gear design is ensuring that the aircraft does not tip over or strike

itself during rotation or landing. Since there is minimal center of gravity excursion, all analysis is

done from the same aircraft location. The major criteria are as follows:

• Longitudinal Tip-over: the angle from the center of gravity to the main gear contact point be

greater than fifteen degrees.

• Lateral Tip-over: two lines are drawn, one (line 1) extending from the main gear to the nose

gear, and another (line 2) perpendicular to the previous line that extends to the center of

gravity. The angle between the ground and line 2 must be less than fifty-five degrees.

• Wing Strike: five degrees between the lowest part of the wing and the main gear.

• Tail Strike: fifteen degrees between the lowest part of the tail and the main gear.

Figure 53 shows that the longitudinal tip-over criteria is met. Figure 54 shows the same for

lateral tip-over. The design of the airplane is such that wing strike and tail strike are of little

concern, with both parts meeting their requirements with ample room to spare.

Lastly, the retraction scheme for the gear can be seen in Figures 55 and 56. The main gear

feature two actuators working in tandem to retract or extend the gear into or out of the fuselage-

wing intersection bulge. The nose gear also features two actuators, and simply folds forward into

the fuselage below the cockpit consoles. These landing gear retraction schemes are valuable for

the overall aircraft designs because they enable easy manufacturability and high reliability, and

help achieve the low acquisition and operating cost goals.

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Figure 53: Longitudinal Tip-over Criteria Figure 54: Lateral Tip-overCriteria

Figure 55: Main Gear Retraction Scheme

Figure 56: Nose Gear Retraction Scheme

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XI. Cabin Design

A large part of the marketing strategy of business jets revolves around the comfort of the cabin.

Companies pride themselves on having tall, wide, spacious cabins outfitted with luxurious leather

seats and large working areas for their clients. With this in mind, a large portion of the design

process was dedicated to making the cabin as large as possible while still maintaining a reasonable

operating empty weight and a long range. After some careful trade studies both in FLOPS and

comparing data to several hundred business jets currently on the market in North America and

Europe, we decided to aim for a cabin height of five feet eight inches and a cabin width of six feet.

These dimensions were designed to be competitive among the eight passenger business jet market.

The cabin for the six passenger is just as spacious, just six feet shorter. As shown in Figure 57, it

is clear that both cabins are very literally heads and shoulders above the competition. Figure 58

shows that both aircraft are also significantly longer than the competition, and can travel just as far

if not farther [21].

A cross section of the H-800 can be seen in Figure 59. The cross section was the same for both

the H-600 and H-800 due to the common fuselages between the variants.

Inside the spacious cabin are plush leather chairs with large executive mahogany tables to work

on. The tables can be raised and lowered as desired, and there is ample leg room between each

seat. One potential cabin layout is shown in Figures 60 and 61. A top view of the H-600 and H-800

can be seen in Figures 62 and 63. Interior baggage storage is available at the aft end of the fuselage

and will be accessible in flight. Additional storage is underneath the empennage. The total payload

capacity is 500 pounds and 60 cubic feet for the H-600 and 1000 pounds and 60 cubic feet for the

H-800. This is more than enough capacity to carry multiple sets of golf clubs, skis, or enough

clothes for a long vacation. The lavatory and galley are located in the cabin section immediately

inside the door and can also be easily accessed in flight. A curtain can be drawn between the

passenger compartment and the galley section for privacy if desired. The entire cabin is lined

with LED lights along the ceiling that can be dimmed from each individual seat for a personalized

experience for each passenger. Large windows provide a significant amount of natural lighting

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Figure 57: Operating Empty Weight versus Cabin Height for Selected FAR 23 Certified BusinessJets [22]

inside the cabin and a spectacular view.

Another important factor in the cabin design was the entertainment and communication system.

On the forward and rear bulkheads there are large screen plasma televisions that are connected to

DirecTV. Inflight Wi-Fi and available above 10,000 feet. Cellular coverage will also be made avail-

able via satellite, ensuring minimal time off the grid in today’s fast-paced corporate environment.

Another element in passenger comfort is the cabin pressure altitude. Most commercial aircraft

pressurize the cabin to the equivalent of 8,000 feet above seal level, with a select few aircraft

pressurizing the cabin as low as 5,000 feet. Higher pressure altitudes result in the passengers

receiving less oxygen during flight, which increases fatigue. Business jet passengers often have

important matters to attend to once they land, and arriving at their destination feeling well-rested is

important. Many business jets currently on the market operate with cabin pressures in the range of

2,000 feet to 3,000 feet at a cruise altitude of 40,000 feet. The current market leader is the SyberJet,

which maintains sea level pressure inside the cabin while cruising at 41,000 feet [23]. Both the

H-600 and H-800 can maintain sea level pressure inside the cabin while cruising at 43,000 feet.

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Figure 58: Maximum Range versus Cabin Length for Selected FAR 23 Certified Business Jets [22]

Figure 59: H-600 and H-800 Cabin Cross Section

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Figure 60: Cabin Interior Rendering

Figure 61: Cabin Interior Rendering

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Figure 62: H-600 Cabin Top View

Figure 63: H-800 Cabin Top View

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XII. Range

As aircraft technology improves, business jets are able to go farther than ever before. Even

small business jets that hold less than 10 people often have ranges well over 2,000 nautical miles.

One of the most travelled business jet routes in the United States is New York to Los Angeles, non-

stop. The target range was 2,500 nautical miles for the H-600 and H-800. Both aircraft, however,

have ranges that well exceed 2,500 nautical miles. With FAR 23 certification, both would have the

longest range in their passenger class by over 200 miles each. This can be the difference between

travelling from the United States to Europe with a single refueling stop or needing to stop twice

[22].

Figure 64: 2,500 Nautical Mile Range Centered On New York

Out of all FAR 23 and FAR 25 certified business jets in the five and six passenger class, the

H-600 has the longest maximum range of them all, and is in the top five percent in its class in range

with the seats full. The H-800 has the longest maximum range out of all FAR 23 certified jets, and

is in the top seventy-five percent out of all FAR 23 and FAR 25 certified jets.

The range of both aircraft will allow them to travel along nine of the top ten most travelled

business jet routes, and seven of the top ten fastest growing routes as well. Within the United

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Figure 65: 2,500 Nautical Mile Range Centered On London

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States, the aircraft are able to travel along all of the most popular vacation routes as well, meaning

you’re a single flight away from hitting the slopes in Aspen after that winter storm dropped several

feet of fresh powder, or a few hours from soaking up the sun in Mexico to get away from the

blustery Chicago winters [24].

Figure 66: Yearly Global Business Jet Flights [25]

Two range maps (one centered on New York and the other on London) in Figures 64 and 65

show the incredible range of the H-600 and H-800. Each map has a range circle of 2,500 nautical

miles highlighted along with a few selected cities to highlight the endless possibilities for flights

on HAMMMR Designs aircraft. Figure 66 shows the tens of thousands of business jet flights that

take place every year.

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XIII. Subsystem Selections

A. Power

The main power generation system for both aircraft will be generators powered by the two

Williams FJ-44 jet turbine engines, shown in Figure 67. Secondary power will be generated by

a Honeywell 36 Auxilary Power Unit (APU). These systems will generate electrical power to op-

erate the internal and external lighting systems, instruments and avionics, passenger comfort sys-

tems, engine starting, flight controls, and de-icing systems, among various other electrical systems

throughout the aircraft.

Figure 67: Williams FJ-44 Jet Turbine Engine [26]

Hydraulic power will be used to operate the retraction and extension of the landing gear, de-

ployment of the flaps, canard movement, and aileron, rudder and elevator movement, as well as

the braking systems. Each of these systems will be triple redundant and will be powered by an

8,000 psi system. Newer aircraft are moving towards using 5,000 or 8,000 psi systems from the

traditional 3,000 psi systems because of the weight saving associated with higher pressure systems.

The aircraft fuel system will consist of two main wing fuel tanks (one in each wing), underbelly

and fuselage tanks, as well as fuel pumps to shift around the location of the fuel to assist with

stability. A diagram of the fuel system can be seen in Figure 68. The wings will hold 1,300 pounds

of fuel each, and the underbelly and fuselage tanks will hold a total of 3,650 pounds of fuel. The

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Figure 68: Fuel System Architecture

tanks, pumps, and crossover tubes are estimated to weigh just under 220 pounds for the entire

system [5].

B. Avionics

Avionics feature heavily in modern aircraft design. The control systems, onboard computers,

communications equipment, and sensor suites make up just some of the electronics that make the

aircraft what it is. The HAMMMR series of aircraft is designed in order to take advantage of

certain advances in avionics in order to offer a high performance, safe flight experience. Critical

to avionics selection was extensive market research to ensure that the series offered low operating

cost and increased capability compared to competitors.

The stability and control system was discussed previously. Other flight oriented avionics in-

clude the autopilot system, associated navigation system, and system monitoring panel in the form

of a glass cockpit. Specifically, a Garmin G5000 flight deck will be installed on the aircraft. High-

resolution displays will allow the flight crew to clearly see even in IFR conditions. Figure 69 shows

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the flight deck layout of a Hawker 400 which has been retrofitted with the G5000 system [27]. Our

cockpit design would be very similar.

Figure 69: Comparable Business Jet Flight Deck Layout [28]

As previously mentioned, the cabin entertainment systems make use of some communications

equipment. Satellite based Internet and television are provided at speeds of up to 50 Mbps through

K_a- band transceivers. But, other antennas and systems are needed for safe operation of the

aircraft. GPS Navigation, accomplished through L-band transceivers, was supplemented by the

Wide Area Augmentation System (WAAS) and VHF Omnidirectional Range (VOR). Additional

VHF radios are also required for aircraft-to-aircraft and aircraft-to-air traffic control communica-

tions. For long haul flights, especially those over water, a HF radio is installed when line-of-sight

communications are unavailable.

Both the H-800 and H-600 are also fitted with weather radars. Newer, smaller radars such

as the Garmin GWX 70 (18-inch aperture, 12.7 pound weight, and 320 nm range) can easily be

incorporated for under $20,000. This feature will allow the aircraft to navigate around bad weather,

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increasing ride comfort. It would also provide the pilot with information to determine if conditions

would be too treacherous for flight. This would allow them to land the aircraft before any danger

was encountered. Figure 70 shows a notional screen display demonstrating the utility of a weather

radar [29].

Figure 70: Weather Radar Display [29]

C. Safety

As an additional safety measure to increase cockpit awareness of nearby aircraft, the HAM-

MMR aircraft series will be outfitted with a second generation Traffic Collision Avoidance Systems

(TCAS II). This system will meet the standards outlined in FAR 135.180 and will be version 7 in

accordance with TSO C-119. Additionally, the system will be tested to ensure it meets the re-

quirements of the International Civil Aviation Organization (ICAO) Standards and Recommended

Practices (SARPS) [30].

In order to combat the buildup of ice during flights, the wings and canard will need either a

de-icing or anti-icing system. In particular, for other canard aircraft the stability of the aircraft

is highly sensitive to icing on the control surfaces. The most commonly used system in business

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aicraft currently is the use of ducts to transport engine bleed air along the leading edge of the wings

to heat the surface and remove any buildup of ice from the surfaces [31]. The aircraft will use this

system for the main wings, and an Electro-Expulsive Separation System (EESS) system for the

canards.

The EESS system uses an electrical current to generate opposing magnetic fields that accelerate

two layers of elastomers apart, breaking any layer of ice up to an inch thick off of the surface of

the outer later. This system uses significantly less power than other thermal de-icing systems and

the ice particles that break off are too small to damage any structures behind the canard, including

the engines [32]. A diagram showing how the EESS system works can be seen in Figure 72

Figure 71: Leading Edge Interior De-Icing Layout [31]

A notional layout of the interior structure of the leading edge of the wing can be seen in Figure

73 and an example layout of a de-icing system from a generic business jet can be seen in Figure

71.

To operate at night, the aircraft will be outfitted with position, anti-collision, taxi, and landing

lights. These lights will be arranged and oeprated in accordance with 14 CFR 91.209. A general

diagram of the required lights for the aircraft can be seen in Figure 74 and a rendering of the H-800

with these lights can be seen in Figure 75.

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Figure 72: Canard De-Icing System UsingElectro-Expulsive Separa-tion System [32]

Figure 73: Business Jet Main Wing De-Icing Sys-tem Layout [31]

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Figure 74: General Air-craft Lighting Layout[33]

Figure 75: Night Rendering of the HAMMMR H-800Showing Lighting Arrangement [34]

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XIV. Alternative Variants

A theme of business jets has been that multiple cabin, and indeed entire aircraft variants, are

produced for different mission sets and customers. Developing additional variants can be beneficial

as it can lower the per-unit development cost and contribute to market awareness for the aircraft.

A. Cargo Variant

The use of a plug for the H-800 easily allows for a cargo variant. Instead of adding an additional

seat section, the plug would feature an over-sized cargo door with dimensions of 5’ x 5’. This,

combined with a stripped down interior for the other three sections of the fuselage, would allow

for 4 standardized North American pallets to be carried up to a maximum weight of 1220 pounds.

B. Electronic/Signals Intelligence

An Electronic/Signals Intelligence (ELINT/SIGINT) could also be made, although it would

require extensive modifications. The Gulfstream G550/650 and Piaggio P180 are just some ex-

amples of business aircraft that have been adapted for these missions. Such an aircraft would be

equipped with computer displays and analysis equipment on one side of the cabin with seats for

the operators. Additional antennas would be required on the outside of the aircraft in bulbs or

suspended. A synthetic aperture radar (SAR) could be mounted in a tub beneath the fuselage for

additional surveillance capability. Figure 76 shows a notional rendering of what this variant might

look like.

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Figure 76: H-800 ELINT/SIGINT Variant (EC-800) with Embedded Antennas (no SAR tub)

XV. Cost Estimation

A. Development and Production Cost

Development and production costs were analyzed throughout the design process using the East-

lake model in order to ensure economic viability. This model analyzed the program cost through

initial operating capability and over a five year production period. It took into account factors such

as percent of composites used, flap complexity, certification type, and was based on the original

DAPCA-IV formulation. The model was adjusted based on the consumer price index to reflect

2017 U.S. Dollar values. A challenge in cost estimation was the fact that both aircraft are of the

same family. To approximate the cost of two separate variants being produced, the engineering cost

was scaled up commensurate to the percent weight difference between variants. Fortunately, tool-

ing costs would be nearly identical since the plug section is common to both aircraft. Additional

analysis was conducted on a pure six person variant to confirm that the results were reasonable.

The RFP requested that a 10 percent profit margin be shown for production rates between 4

and 10 aircraft per month. Reports indicated that the business jet market for this class of aircraft is

projected to take around 300 deliveries a year into 2025. Table XXXI shows the necessary market

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penetration percentage to meet production levels and the associated unit cost.

Table XXXI: Market Penetration Required for given Aircraft Production Level and AssociatedUnit Cost

Market Penetration(%)

Production Rate(Aircraft Per Month)

Aircraft Unit Cost(FY 2017 Million USD)

4 1 20.18 2 12.7

12 3 9.916 4 8.420 5 7.424 6 6.728 7 6.232 8 5.836 9 5.540 10 5.2

Historically, new business jet aircraft have been projected to achieve around 25% market pen-

etration. This is further backed up by analysis of other business products which generally capture

between 10% and 40% of the market.

If this were the case, then the aircraft could be produced at a rate of 6 aircraft per month (360

over the five-year production period) and sold for a price of $7.4 million for a 10% profit margin. It

is likely, however, that the H-800 would be priced at $7.5 million, with the H-600 at $6.5 million. A

key variable here is which variant would sell more. Sales of the H-800 in a faster time to break-even

since has an increase in profit of $100,000 over the H-600. It is possible, however, that the H-800

and H-600 would not compete for sales but instead be complementary with each taking different

parts of the market. Figure 77 shows the effect of price (of the H-800 alone) on time to break-even,

with the obvious caveat that higher prices would likely result in lower sales. Conversely, however,

a lower price might result in a greater long-term profit margin, but such analysis is beyond the

scope of this report.

At the 6 aircraft per month production rate, the Eastlake model yielded the cost breakdown

shown in Figure 78. Of this, approximately 35% is non-recurring since engineering, tooling, and

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Figure 77: Revenue as Function of Aircraft Cost

flight test are are generally a one time expense. It should be noted, however, that if the aircraft’s

production had to be massively scaled up, additional tooling and engineering support could be

required.

B. Operating Cost

A critical objective of the design process was minimizing operating cost. Specifically, by

reducing fuel burn, variations in the price of fuel would have less of an impact on flight operations.

The operating costs shown in Figure 79 are for an aircraft flying approximately 1000 hours a year.

While this is a high estimate for an individual operator, the aim is to show that the HAMMMR

series makes a strong business case for charter aircraft operators. Additionally, any reduction in

the price of operation is also beneficial to an individual operator. Figure 80 shows the variation in

operating cost as a function of number of hours flown. As the figure shows, there are substantial

operating cost benefits in increasing the flight hours per year to around 1000. After that, however,

additional flight hours per year have less of an effect on the operating cost.

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Figure 78: Acquisition Cost Breakdown for 360 Aircraft

A possible business case for the aircraft could be made based on this information. If flown

1,000 hours per year, the cost per flight hour works out to be $1,250. Given that the typical block

time is 6 hours, the total price for a flight is $7,500. If sold for $12,000 ($2,000 per hour), a 60$

profit but not unreasonable for an eight person transcontinental flight, then a company will make

approximately $750,000 per year per aircraft. At this rate, a company will reach the break-even

point after ten years. It is also likely that a company doing a bulk order could negotiate a lower

selling price for the aircraft. It should also be noted that this cost per flight hour, $2,000, which

again represents 60% profit, is below the industry standard cost [35].

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Figure 79: Operating Cost Breakdown at 1000 Hours per Year

Figure 80: Variation in Cost per Flight Four with Number of Hours Flown per Year

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XVI. Conclusion

This technical proposal outlined the design and development of a two-member light business

jet aircraft family with capacities of six and eight passengers. The goal of this project was to

introduce new technology and increased capability with a low operating cost in order to meet

current and future demands.

The eight-passenger variant was used as the primary design driver given its higher payload

and resulting performance impact. Multiple Class I concepts were considered; these included

conventional and unconventional component configurations across a variety of weight classes.

Ultimately, a canard-equipped aircraft with a maximum gross takeoff weight of roughly 17,000

lbs was selected. Rigorous technical analysis was performed to validate this design choice using

several commercial software packages for computational fluid dynamics, finite element analysis,

trade-study-guided optimization, and computer-aided design and simulation. In addition, several

custom software applications were developed in order to optimize the wing and canard size, loca-

tion, and stability characteristics. Due to the highly-coupled nature of the design, all analyses were

extensively iterated.

To guarantee the creation of a feasible light business jet family, both the six- and eight-passenger

variants were expected to share at least 70% parts commonality. Consequently, the six-passenger

aircraft demanded extensive design consideration to meet this requirement effectively. Ultimately,

a simple fuselage plug, six feet in length, was chosen as main differentiating element between the

two aircraft. While this choice complicated the center of gravity and stability characteristics to

achieve the desired performance, it significantly reduced the overall design complexity and created

a net reduction in development and acquisition costs. Since the wing and canard were sized to

meet the eight-person requirements, they are non-optimal for the six-person design mission. This

does, however, allow the six-person to have a greater range and reduced cost due to increased parts

commonality and shared tooling. Performance and cost analyses validated this decision.

Financially, these clean-sheet aircraft represent a unique and lucrative opportunity for a new

or existing manufacturer in the light business jet market. Based on extensive cost modeling and

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analysis, the current-year sale price of the eight-passenger aircraft was found to be US $7.5 million

while the current-year sale price of the six-person variant was found to be US $6.5 million. These

prices include a 15% profit margin for the manufacturer at a production rate of 6 aircraft per month.

The emphasis on feasibility and technology readiness throughout the design process results in high

confidence that these aircraft would meet a target entry-into-service date of 2020 and 2022 for the

six- and eight-passenger aircraft, respectively.

In light of the growing market potential and high financial upside, this light business jet aircraft

family shows substantial promise and constitutes a worthy business venture. Through the strategic

use of advanced yet proven technologies, commercial-off-the-shelf subsystems, and modern manu-

facturing practices, the HAMMMR Designs family of aircraft successfully delivers a sophisticated

aesthetic while accomplishing an impressive performance envelope and guaranteeing maximum

value for both customers and shareholders. It is believed that the HAMMMR series will offer a

compelling choice for both individual and multi-aircraft buyers. Go Jackets, and Fly HAMMMR!

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XVII. Appendix

Figure 81: HAMMMR H-800 with Nighttime Operating Lights

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