Design of a Light Business Jet Family David C

Design of a Light Business Jet Family David C. Alman Andrew R. M. Hoeft Terry H. Ma AIAA : 498858 AIAA : 494351 AIAA : 820228

Cameron B. McMillan Jagadeesh Movva Christopher L. Rolince AIAA : 486025 AIAA : 738175 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 Page i 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

Page ii 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.

Page iii 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, location, 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

Page iv 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 manufacturing 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.

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

Page vi Contents

I Acknowledgements i

II Executive Summary iv

III Requirements 1

IV Business Jet Market Analysis 2

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

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

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

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

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

X Landing Gear Placement and Design 58

XI Cabin Design 61

Page vii XII Range 66

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

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

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

XVI Conclusion 82

XVII Appendix 84

List of Figures 1 Scaled Three-View Image of the HAMMMR Designs H-800 ...... vi 2 FAA Data on Market Growth ...... 2 3 Standard T-Tail, Cruciform Tail, and V-Tail Configurations ...... 9 4 Selected Component Configuration ...... 10 5 NBAA IFR Range Profile ...... 11 6 Energy based constraint sizing ...... 13 7 Historical Aircraft Weight Regression (pictures not to scale) ...... 15 8 Trade Study of Cruise Conditions Effect on Gross Weight ...... 16 9 Trade Study of Cruise Conditions Effect on Design Mission Block Time . . . . . 17 10 FLOPS Output Drag Polar at Mach 0.8, 43k ft ...... 18 11 H-800 Payload Range Diagram ...... 19 12 H-600 Payload Range Diagram ...... 19 13 Location of Aircraft Datum ...... 22 14 Component Weight Layout of 6 Passenger Variant ...... 22 15 Component Weight Layout of 8 Passenger Variant ...... 23 16 Center of Gravity Excursion Diagram for 8 Passenger Variant ...... 24 17 Center of Gravity Excursion Diagram for 8 Passenger Variant ...... 24 18 Section Lift Coefficient ...... 26 19 Section Drag Coefficient ...... 26 20 Surface Growth Rate of Wing Meshes ...... 27 21 Boundary Layer Mesh Refinement ...... 27 22 Relative Mach Number of Canard ...... 28 23 Relative Mach Number of Main Wing ...... 28 24 Planforms of Wing Surfaces. Controls are scaled correctly relative to surface, but surfaces are not to scale ...... 29

Page viii 25 Locations Considered for Wing and Canard Placement ...... 32 26 Canard Moment Analysis ...... 33 27 Depiction of Aerodynamic Coupling with Canard ...... 34 28 Aerodynamic Benefit of Canard and Wing Coupling ...... 34 29 Stall Strip for Canard ...... 35 30 Configuration Optimizer Execution Process Flow ...... 37 31 Potential 6 and 8 Person Static Margin Shifts with Maximum Parts Commonality 38 32 Short-period Flying Qualities ...... 42 33 Longitudinal SCAS Block Diagram ...... 43 34 Block diagram of Lateral SCAS ...... 45 35 Overview of Aircraft Structural Layout ...... 46 36 Structural Configuration and Material Selection for Aircraft ...... 47 37 Fuselage Construction ...... 48 38 Fuselage Panel Layup ...... 49 39 Simplified Fuselage Structural Model and Mesh ...... 49 40 IRF Plot for Pressurized Composite Fuselage ...... 49 41 Wing-Fuselage Attachment ...... 50 42 Wing Spar to Fuselage Interface ...... 50 43 Wing Structural Model and Mesh ...... 51 44 Wing Internal Structure Mesh ...... 52 45 Pressure Load Application on Bottom Surface of Wing ...... 53 46 Shape of Pressure Distribution ...... 53 47 Wing Deformation under Load ...... 54 48 Tail Structural Layout ...... 55 49 Tail Attachment to Empennage ...... 55 50 Canard Articulation ...... 55 51 V-N Diagram for 6 Passenger Variant ...... 56 52 V-N Diagram for 8 Passenger Variant ...... 57 53 Longitudinal Tip-over Criteria ...... 60 54 Lateral Tip-over Criteria ...... 60 55 Main Gear Retraction Scheme ...... 60 56 Nose Gear Retraction Scheme ...... 60 57 Operating Empty Weight versus Cabin Height for Selected FAR 23 Certified Business Jets ...... 62 58 Maximum Range versus Cabin Length for Selected FAR 23 Certified Business Jets 63 59 H-600 and H-800 Cabin Cross Section ...... 63 60 Cabin Interior Rendering ...... 64 61 Cabin Interior Rendering ...... 64 62 H-600 Cabin Top View ...... 65 63 H-800 Cabin Top View ...... 65 64 2,500 Nautical Mile Range Centered On New York ...... 66 65 2,500 Nautical Mile Range Centered On London ...... 67 66 Yearly Global Business Jet Flights ...... 68 67 Williams FJ-44 Jet Turbine Engine [26] ...... 69 68 Fuel System Architecture ...... 70 Page ix 69 Comparable Business Jet Flight Deck Layout ...... 71 70 Weather Radar Display ...... 72 71 Leading Edge Interior De-Icing Layout ...... 73 72 Canard De-Icing System Using Electro-Expulsive Separation System ...... 74 73 Business Jet Main Wing De-Icing System Layout ...... 74 74 General Aircraft Lighting Layout ...... 75 75 Night Rendering of the HAMMMR H-800 Showing Lighting Arrangement . . . 75 76 H-800 ELINT/SIGINT Variant (EC-800) with Embedded Antennas (no SAR tub) 77 77 Revenue as Function of Aircraft Cost ...... 79 78 Acquisition Cost Breakdown for 360 Aircraft ...... 80 79 Operating Cost Breakdown at 1000 Hours per Year ...... 81 80 Variation in Cost per Flight Four with Number of Hours Flown per Year . . . . . 81 81 HAMMMR H-800 with Nighttime Operating Lights ...... 84

List of Tables I Matrix of Requirement Compliance ...... 1 II Fuselage Configuration Figure of Merit ...... 5 III Wing Configuration Figure of Merit ...... 6 IV Landing Gear Configuration Figure of Merit ...... 6 V Engine Number Figure of Merit ...... 8 VI Engine Failure Probability ...... 8 VII Tail Configuration Figure of Merit ...... 9 VIII Mission Segments for NBAA IFR Profile...... 12 IX Constraint Sizing Outputs ...... 14 X 8 Person Sizing Mission: Crew and Payload Weights ...... 14 XI 6 Person Sizing Mission: Crew and Payload Weights ...... 15 XII State of Design ...... 18 XIII Component Weights ...... 21 XIV Comparison of Airfoil Series ...... 25 XV Characteristics of Main Wing ...... 30 XVI Characteristics of Vertical Tail ...... 30 XVII Winglet Design Effectiveness [10] ...... 31 XVIII Characteristics of Canard ...... 35 XIX Configuration Optimizer Parameter Sweep ...... 36 XX Configuration Optimization Priorities ...... 39 XXI Sensitivity Study on Level of Commonality Versus Optimal Configuration Point for H-600 and H-800 ...... 39 XXII Final Configuration Placement ...... 40 XXIII Final Configuration Static Margins...... 40 XXIV Longitudinal Flying Qualities ...... 41 XXV Natural Longitudinal Flying Qualities ...... 42 XXVI Lateral Flying Qualities ...... 44 XXVII Natural Lateral Flying Qualities ...... 44 XXVIII Smeared Properties of 24 ply Carbon Fiber Laminates [18] ...... 51

Page x XXIX Landing Gear Loads ...... 58 XXX Landing Gear Characteristics ...... 59 XXXI Market Penetration Required for given Aircraft Production Level and Associ- ated Unit Cost ...... 78

Nomenclature List of Acronyms ACP ANSYS Composite PrePost AIAA The American Institute of Aeronautics and Astronautics AVL Athena Vortex Lattice CD Drag coefficient CFD Computational fluid dynamics CFR Code of Federal Regulations DAPCA Development and Production Costs for Aircraft ELINT Electronics intelligence FAA Federal Aviation Administration FLOPS Flight Optimization System HF High Frequency ICAO International Civil Aviation Organization IRF Inverse Reserve Factor LED Light-emitting diode MGTOW Maximum gross takeoff weight NACA National Advisory Committee for Aeronautics NASA National Aeronautics and Space Administration NBAA National Business Aviation Association RFP Request for proposal SAR Synthetic aperture radar SARPS Standards and Recommended Practices SAS Stability augmentation system SIGINT Signals intelligence SST Menter’s Shear Stress Transport turbulence model TCAS Traffic Collision Avoidance System TSFC Thrust-specific fuel consumption USD United States Dollar VHF Very-High Frequency VOR Very-High Frequency Omnidirectional Range WAAS Wide Area Augmentation System List of Symbols c¯ Mean Chord Length Γw Wing Dihedral Angle Λc/4 Quarter-Chord Sweep Angle λw Wing Taper Ratio AR Aspect Ratio

Page xi cf Flap Chord CL,max Maximum Lift Coefficient CL Lift coefficient cr Rudder Chord CD0 Zero-Lift Drag Coefficient CG Center of Gravity e Oswald efficiency number L/D Lift-to-drag ratio LE Leading Edge S Main Wing Planform Area Sa Aileron Planform Area Sc Canard Planform Area Sf Flap Planform Area Sr Rudder Planform Area Vv Vertical Tail Volume Ratio Xcg Center of Gravity X-location XNP Neutral Point X-location List of Units fpm feet per minute ft feet in inches KIAS knots indicated airspeed nm nautical miles psf pounds per square foot sq ft square feet

Page xii 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.B Rate 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.B Takeoff 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.C Minimum 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 XIII Baggage Capacity 500 pounds / 30 ft3 1,000 pounds / 60 ft3 500lbs / 60 ft3 1000lbs / 60 ft3 XII Passenger Capacity 6 pax 8 pax 6 pax 8 pax XII

Page 1 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,

Page 2 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 competitive 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 customers 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 business 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 opportunities 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 acquisition 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 characteristics, or ﬁgures of merit, of the aircraft. The weighting factors, based on the stakeholder require-

Page 4 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 Fuselage Manufacturability 3.80 4.83 3.00 2.67 Weight Savings 3.80 3.17 3.00 3.33 Aerodynamic Performance 3.80 3.17 3.50 5.00 Passenger 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 conventional, 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 manufacturing 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.

Page 5 B. Wing Conﬁguration

Table III: Wing Conﬁguration Figure of Merit

Wing Weighting − − − − − Low Wing High Wing Mid Wing Stability 4.20 2.80 4.40 3.20 Landing Gear 3.20 4.00 2.40 3.20 Maintenance 3.80 4.40 2.60 3.20 Passenger Comfort 3.20 3.40 3.40 2.80 Aerodynamic 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 Quadruple Component 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

Page 6 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.

Page 7 F. Engine Number

Table V: Engine Number Figure of Merit

Engine Number Weighting − − − − 2 Engines 3 Engines Reliability 3.80 3.33 4.00 Maintenance Costs 3.40 4.33 2.17 Fuel Efﬁciency 3.60 4.40 2.20 Weight 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 Engines 2 2 Engines 2Pef Pef N/A 2 3 3 Engines 3Pef 3Pef Pef

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.

Page 8 G. Tail Conﬁguration

Table VII: Tail Conﬁguration Figure of Merit

Tail Weighting − − − − − T-Tail V-Tail Cruciform Canard Aerodynamic 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 conventional 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, tricycle 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 profile 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 - 43,000 Cruise 2180 43,000 Descent 110 43,000 - 0 Land - - 100 nm Alternate Divert Climb 14 0 - 18,000 Cruise 44.2 18,000 Descent 41.8 18,000 - 0 Land - -

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

Page 12 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

TSL/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

Page 14 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 considerations. 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

Page 16 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 competitive 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 significant 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.

Page 18 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 0.17 Anti-Icing 75 38.3 1.7 32.3 0.17 Auxiliary Power 321 42.5 1.7 36.5 1.7 Avionics 408 41 -0.5 35 -0.5 Canard 141 5.3 -1 5.3 -1 Cargo Weight 250 41.7 -1.17 35.7 -1.17 Crew and Baggage 225 6.25 0.33 6.25 0.33 Electrical 953 40 0 34.0 0 Engine Oil 31 45.3 4.16 39.3 4.16 Engines 1019 45.3 4.16 39.3 4.16 Fuel System 218 33.1 -1.5 27.1 -1.5 Furnishings 1717 / 1287.75 30 0 22 0 Fuselage 1851 / 1150 28.8 0 25.8 0 Fuselage Fuel 4370 / 4060 33.1 -2.5 27.1 -2.5 Hydraulics 227 38.3 -1.7 32.3 -1.7 Instruments 181 3.75 -2.2 3.75 -2.2 Main Gear 355.6 34.7 -2.2 28.7 -2.2 Main Wing 1399 34.7 -2.2 28.7 -2.2 Nose Gear 152.4 7.5 -2.17 7.5 -2.17 Passenger Baggage 120 / 160 41.7 -1.17 35.7 -1.17 Passenger Service 125 33.3 -1 27.3 -1 Passengers 540 / 720 24.17 -1 18.17 -1 Surface Controls 333 34.7 -1.08 28.7 -1.08 Unusable Fuel 264 34.1 -2.5 28.1 -2.5 Vertical Tail 164 41.7 6.25 35.7 6.25 Wing 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.

Page 23 Figure 16: Center of Gravity Excursion Figure 17: Center of Gravity Excursion Diagram for 8 Passenger Variant Diagram 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 considerations 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 Good Stall Characteristics Low Maximum Lift Coefficient General Aviation 4-Series Small Deviation of Center of Pressure Higher Drag Coefficient Horizontal Tails Not Sensitive to Surface Roughness Large Pitching Moment High Maximum Lift Coefficient Poor Stall Behaviour General Aviation Low Pitching Moment Relatively High Drag Piston Aircraft 5-Series Not Sensitive to Surface Roughness Commuters Business Jets High Maximum Lift Coefficient High Non-Optimal Drag Piston Fighters Low Optimal Drag High Pitching Moment Business Jets 6-Series Optimized for High Speed Poor Stall Behavior Jet Trainers Susceptible 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 conditions 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 Number Figure 23: Relative Mach Number of Main of Canard Wing

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 considered 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 respectively. 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 design, 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 aircraft. 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, but surfaces are not to scale

The wing dihedral, Γw, was determined through both ground clearance and stability considerations. 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 incidence 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 Value Wing 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 0.105 λW 0.35

Table XVI: Characteristics of Vertical Tail

Parameter Value Parameter Value Wing Area 81 ft2 Aspect Ratio 5.54 Airfoil NACA 0009 Λc/4 45° Root Chord 9.0 ft Vv 0.06 Half 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 configurations. These included configurations with no winglet, a blended winglet, a scimitar winglet. The results of this comparison are seen below in Table XVII below.

Page 30 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

Page 31 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.

x xcg S taticMargin = NP − (1) c¯ c¯

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

Page 32 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 coefficient 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 Page 33 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 Aerodynamic Figure 28: Aerodynamic Beneﬁt of Ca- Coupling with Canard [5] 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.

Page 34 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 Value Wing 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

Page 35 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 9 Engine X-location (ft) 38.17 45.83 9 Engine Z-location (ft) 1.25 5.0 9 Fuselage Fuel x-location (ft.) 31.25 43.33 9 Canard 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 important 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.

Page 37 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.

Page 38 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 for H-600 and H-800

Variant Wet Static Margin (%) Dry Static Margin (%) Static Margin Shift (%) Only Plug H-600 6.6 3.3 3.2 H-800 −3.2 −1.3 −1.9 Plug + Fuselage Fuel Change H-600 0.9 2.9 −2.0 H-800 0.1 0.2 0.1 Plug + Fuselage Fuel + Canard Area H-600 0.0 0.0 0.0 H-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 Page 39 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 33.1 H-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 −2.0 H-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 Page 40 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 performance. 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 Mode Level 1 ζ > 0.04 Level 2 ζ > 0 Level 3 T2 > 55s Short-period Mode ζsp ωn, sp min. max. min. max. Level 1 0.4 1.3 2.4 3.8 Level 2 0.25 1.3 1.8 6.3 Level 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].

   Xu Xw −   m m 0 gcosθ0     Z Z Zq+mU0 −mgsinθ   u w 0  A =  m−Zω˙ m−Zω˙ m−Zω˙ m−Zω˙  (2) lon   I−1[M + Z Γ] I−1[M + Z Γ] I−1[M + (Z + mU )Γ] −I−1mgsinθ Γ  yy u u yy w w yy q q 0 yy 0     0 0 1 0 

Page 41 Figure 32: Short-period Flying Qualities [16]

 X   Xδe δp   m m     Zδ Zδp   e  B  m−Zω˙ m−Zω˙  lon =   (3) I−1[M + Z Γ] I−1[M + Z Γ]  yy δe δe yy δe δe     0 0 

0 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 Qualities Mode Level 1 Requirement Result Analysis Phugoid ζ > 0.04 ζ = 0.0087 Unsatisfactory 0.4 < ζ < 1.3 ζ = 0.36 Unsatisfactory Short-Period 2.4 < ω < 3.8 ω = 2.99 Satisfactory

Page 42 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.

Page 43 Table XXVI: Lateral Flying Qualities

Roll Mode Time Constant Level 1 20s Level 2 12s Level 3 4s Spiral Mode Minimum Doubling Time Level 1 12s Level 2 8s Level 3 4s Dutch Roll Mode ζmin ωn,min ζ ∗ ωn Level 1 0.19 1 0.35 Level 2 0.02 0.4 0.05 Level 3 0.02 0.04 No Limit

Once again, it was necessary to populate a state space model of the aircraft with the appropriate 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 Qualities Mode Level 1 Requirement Result Analysis Roll t < 20s 7.38s Unsatisfactory Spiral t < 12s 12.63s Satisfactory ζ > 0.19 0.0038 Unsatisfactory Dutch Roll ωn > 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.

Page 44 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 elements. 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 layers, 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 (finite 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

Page 46 Figure 36: Structural Conﬁguration and Material Selection for Aircraft (Target Weight Reduction Values 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 geometry 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

Page 48 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 ksi Poisson’s Ratio 0.305 Shear 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:

√ 2W n L2 − z2 q z TO ( ) = 2 (6) piL ∗ 2 ∗ [(Cr − Ct)/b + Cr]

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

Page 56 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.

Page 58 Table XXX: Landing Gear Characteristics

- Front Gear Main Gear Number of Tires 1 2 Tire 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.

Page 59 Figure 53: Longitudinal Tip-over Criteria Figure 54: Lateral Tip-over Criteria

Figure 55: Main Gear Retraction Scheme

Figure 56: Nose Gear Retraction Scheme

Page 60 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

Page 61 Figure 57: Operating Empty Weight versus Cabin Height for Selected FAR 23 Certified Business Jets [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 available 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. Page 62 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

Page 63 Figure 60: Cabin Interior Rendering

Figure 61: Cabin Interior Rendering

Page 64 Figure 62: H-600 Cabin Top View

Figure 63: H-800 Cabin Top View

Page 65 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

Page 66 Figure 65: 2,500 Nautical Mile Range Centered On London

Page 67 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.

Page 68 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 operate the internal and external lighting systems, instruments and avionics, passenger comfort systems, 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, deployment 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 Page 69 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 include 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

Page 70 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 communications. 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,

Page 71 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 requirements 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

Page 72 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.

Page 73 Figure 72: Canard De- Figure 73: Business Jet Main Wing De-Icing Sys- Icing System Using tem Layout [31] Electro-Expulsive Separa- tion System [32]

Page 74 Figure 74: General Air- Figure 75: Night Rendering of the HAMMMR H-800 craft Lighting Layout Showing Lighting Arrangement [34] [33]

Page 75 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.

Page 76 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, tooling 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

Page 77 penetration percentage to meet production levels and the associated unit cost.

Table XXXI: Market Penetration Required for given Aircraft Production Level and Associated Unit Cost

Market Penetration Production Rate Aircraft Unit Cost (%) (Aircraft Per Month) (FY 2017 Million USD)

4 1 20.1 8 2 12.7 12 3 9.9 16 4 8.4 20 5 7.4 24 6 6.7 28 7 6.2 32 8 5.8 36 9 5.5 40 10 5.2

Historically, new business jet aircraft have been projected to achieve around 25% market penetration. 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

Page 78 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.

Page 79 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].

Page 80 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

Page 81 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, location, 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

Page 82 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 manufacturing 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!

Page 83 XVII. Appendix

Figure 81: HAMMMR H-800 with Nighttime Operating Lights

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