PegasusAerospaceP65

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The Pegasus Aerospace Team
Nicholas Robson
Georgia Tech BSAE ‘14
Member #: 487875
Kieffer Milligan
Collin Strassburger
Member #: 463147
Airth Burtman
Christian
Rasmussen
Andrew Norris
Josh Rogers

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Executive Summary
The Pegasus Aerospace Pegasus 65 (P-65) is a rear-engine box-wing regional turboprop that
addresses the need demonstrated by the aviation industry for a 75 passenger future turboprop
aircraft with 65% reduction in fuel consumption compared to existing jet aircraft. This is in
response to a Request for Proposal (RFP) by the
American Institute of Aeronautics and
Astronautics (AIAA)2.
Fuel Reductions
The P-65 aircraft uses 5,399 LBS of Jet-A fuel for
its 400 nautical miles (nmi) economic design
mission. When compared to existing competing
regional jets (CRJ700), this reduces fuel
consumption by 68.8% for the same mission –
exceeding the RFP requirement.
New Technologies
An Ultra High Bypass Turbofan utilizing a PW-150
derivative core powers the P-65 aircraft, and
contributes to a proven 30% reduction in fuel
consumption. Natural laminar flow (NLF)
aerodynamics are utilized for the nosecone
and airfoils, which contribute to a 24.4%
reduction in zero lift drag during cruise as
predicted by Class I and II drag polar approximations, and vortex lattice computations. Carbon
composite materials used in the fuselage contribute to a 15% savings in fuselage weight.
Additionally, the innovative box wing design reduces induced drag and increases effective
span which leads to an increase in Oswald’s efficiency factor from 0.91 to 1.47, when compared
Aircraft Summary:
Configuration: Rear-Engine Box wing
Design Mission: 400NMI Economic Mission
Payload: 74 Passengers, 4 Crew
Weight: WTO – 45,250 LBS, WOE – 39,850 LBS
Design Cruise Speed: Mach .68
Operating Altitude: 31,000 Ft.
Fuel Consumption: WFuel - 5,399LBS @
400NMI Economic Mission
Fuel Burn Reduction: 68.8% fuel burn
reduction compared to existing regional jet
for same mission

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to conventional aircraft designs. The box wing’s braced structure also allows for a 40% reduction
in necessary wing structure and weight. In all, the P-65 Aircraft reduces weight by 70.1% when
compared to existing regional aircraft.
Jet-Like Experience
The blade design of the Ultra-High Bypass Ratio engine allows the P-65 to cruise at Mach 0.68
while maintaining the efficiency benefits of traditional turboprop. The rear-engine configuration
minimizes acoustic propagation of engine noise to the passenger compartment, and active and
passive noise cancellation combined with a plenum environmental control air distribution
network contribute to a passenger cabin that is 35.3% quieter than the industry standard.
Cost and Environmental Friendliness
The material selection and manufacturing processes selected for the P-65 allow it to have a
competitively priced flyaway cost of $24.4M US Dollars (USD) per aircraft based on a 400 aircraft
production run. Reduced fuel consumption makes the P-65 an extremely Environmentally
Responsible Aircraft (ERA), reducing carbon emissions by 69.2% while simultaneously reducing
Operation and Maintenance (O&M) costs for airlines. O&M costs (7.8¢ cost per seat mile) for the
P-65 are reduced by 11.4% and 30.3% for existing regional jet and turboprop respectively.

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Table of Contents
The Pegasus Aerospace Team.....................................................................................................ii 
Executive Summary ...................................................................................................................... iii 
Fuel Reductions......................................................................................................................... iii 
New Technologies .................................................................................................................... iii 
Jet-Like Experience...................................................................................................................iv 
Cost and Environmental Friendliness......................................................................................iv 
Table of Contents..........................................................................................................................vi 
List of Figures .................................................................................................................................vii 
List of Tables ...................................................................................................................................ix 
Nomenclature ................................................................................................................................x 
Introduction ................................................................................................................................... 1 
Regional Jet Outlook.................................................................................................................... 1 
Market Analysis ......................................................................................................................... 2 
Stakeholder Analysis................................................................................................................. 5 
Existing Designs.............................................................................................................................. 6 
Configuration Selection ............................................................................................................... 7 
Figures of Merit .......................................................................................................................... 7 
Configuration Selection ......................................................................................................... 15 
Technologies................................................................................................................................ 16 
Box Wing .................................................................................................................................. 16 
Propfan .................................................................................................................................... 17 
Composite Structures ............................................................................................................. 17 
Natural Laminar Flow ............................................................................................................. 17 
Digital fly-by-wire .................................................................................................................... 18 
Solid Oxide Fuel Cell Auxiliary Power Unit ............................................................................ 18 
Riblets ....................................................................................................................................... 18 
Spyroid Wingtip ....................................................................................................................... 18 
Circulation Control Wing ....................................................................................................... 18 
Technology Compatibility ..................................................................................................... 19 
Weight Sizing................................................................................................................................ 20 
Mission Profile........................................................................................................................... 20 
Weight Estimation ................................................................................................................... 20 
Payload Weight ...................................................................................................................... 21 
Mission Fuel Weight................................................................................................................. 21 
Gross Takeoff and Empty Weights ........................................................................................ 22 
Drag Polar (Class I Method) .................................................................................................. 23 
Drag Polar (Class II Method).................................................................................................. 25 
Constraint Sizing .......................................................................................................................... 26 
K Determination ...................................................................................................................... 27 
Constraint Sizing Results ......................................................................................................... 29 
Wing Design................................................................................................................................. 31 
Planform Selection.................................................................................................................. 31 
Airfoil Selection........................................................................................................................ 34 
Empennage Sizing.................................................................................................................. 34 
High Lift Devices...................................................................................................................... 35 
Control Surfaces ..................................................................................................................... 36 
Stability and Control............................................................................................................... 37 
Weight and Balance .................................................................................................................. 39 
Design Point Selection Iterative Process................................................................................... 41

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Fuselage....................................................................................................................................... 46 
Sizing......................................................................................................................................... 47 
Noise Reduction...................................................................................................................... 49 
Doors ........................................................................................................................................ 49 
Structure................................................................................................................................... 49 
Flight Deck Design .................................................................................................................. 52 
Propulsion..................................................................................................................................... 54 
Core Selection ........................................................................................................................ 54 
Number of Blades ................................................................................................................... 54 
Tip Speed ................................................................................................................................. 55 
Blade and Disk Power Loading ............................................................................................. 55 
Blade Sweep ........................................................................................................................... 56 
Blade Twist and Taper ............................................................................................................ 57 
Blade Materials ....................................................................................................................... 57 
Inlet........................................................................................................................................... 57 
Results and Summary ............................................................................................................. 58 
Landing Gear .............................................................................................................................. 59 
Configuration .......................................................................................................................... 59 
Tip-Over and Clearance........................................................................................................ 59 
Tires ........................................................................................................................................... 62 
Retraction Kinematics ............................................................................................................ 62 
Structural Consideration............................................................................................................. 64 
Materials .................................................................................................................................. 64 
V-n Diagram............................................................................................................................ 64 
Environmental Impact................................................................................................................ 65 
Systems......................................................................................................................................... 66 
ECS ........................................................................................................................................... 66 
Fuel System .............................................................................................................................. 71 
Electrical System ..................................................................................................................... 73 
Cost Analysis ................................................................................................................................ 74 
Flyaway Cost........................................................................................................................... 74 
O&M Costs............................................................................................................................... 77 
Marketing................................................................................................................................. 79 
References................................................................................................................................... 84 
List of Figures
Figure 1: Air Traffic RPK Growth Through 20245 .......................................................................... 2 
Figure 2: Air Traffic by Flow5.......................................................................................................... 3 
Figure 3: Regional Aircraft Proportion of World Fleets5 ............................................................. 3 
Figure 4: Average Gasoline Prices, 1991-201258 ........................................................................ 4 
Figure 5: Global LCC Growth11 .................................................................................................... 4 
Figure 6: Stakeholder Analysis...................................................................................................... 5 
Figure 7: Q40025 and CRJ70024..................................................................................................... 6 
Figure 9: Box Wing Configuration23 ........................................................................................... 10 
Figure 10: Turboprop21 and Propfan Engines28 ........................................................................ 12 
Figure 11: Propfan, Turboprop and Turbofan Propulsive Efficiency Plots31........................... 13 
Figure 12: Preliminary Design Configuration ............................................................................ 15 
Figure 14: Mission Profile ............................................................................................................. 20 
Figure 15: Weight Regression for Comparable Aircraft29 ....................................................... 23 
Figure 17: Comparison of Class I and Class II Drag Polar – Cruise......................................... 26

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Figure 18: Energy Based Constraint Sizing ................................................................................ 31 
Figure 19: Altitude vs. Fuel Burn for Multiple Mach at AR = 12 ............................................... 32 
Figure 20: Altitude vs. Fuel Burn for Multiple AR at M = 0.68 ................................................... 33 
Figure 21: Fuel Weight vs. Sweep Angle for Multiple AR......................................................... 33 
Figure 22: Final Wing Configuration .......................................................................................... 33 
Figure 23: Control Surface Diagram ......................................................................................... 37 
Figure 24: Longitudinal SAS Control39........................................................................................ 39 
Figure 25: SAS Response Properties39 ........................................................................................ 39 
Figure 26: CG Excursion Diagram.............................................................................................. 41 
Figure 27: Iterative Process Workflow........................................................................................ 42 
Figure 28: Climb Speed vs. Altitude .......................................................................................... 43 
Figure 29: Tail Forward on CG Travel for Multiple c (f/a)........................................................ 44 
Figure 30: Differential Alpha on CG for Multiple c (f/a) ......................................................... 44 
Figure 31: Sweep (f/a) on CG Travel for Multiple c (f/a)........................................................ 44 
Figure 32. P-65 Cabin Cross-Section ......................................................................................... 47 
Figure 33. Overall Fuselage Dimensions ................................................................................... 48 
Figure 34. Fuselage Interor Configuration ................................................................................ 48 
Figure 35. Fuselage Structure Model......................................................................................... 50 
Figure 36. Sheet-Stringer Approximation9................................................................................. 51 
Figure 37. Bay Structure (Dimensions in Inches)....................................................................... 52 
Figure 38. P-65 Cockpit Visibility................................................................................................. 53 
Figure 39. P-65 Cockpit17 ............................................................................................................ 53 
Figure 40: Blade Diameter vs. Blade loading........................................................................... 55 
Figure 41: Spin Rate vs. Power Loading per blade ................................................................. 56 
Figure 42: Final Engine Design ................................................................................................... 59 
Figure 43: Longitudinal Tip-Over Criterion Check47 ................................................................. 60 
Figure 44: Lateral Tip-Over Criterion47 ....................................................................................... 61 
Figure 45: Lateral Tip-Over Criterion Check ............................................................................. 61 
Figure 46: Longitudinal Ground Clearance Check................................................................. 61 
Figure 47: Lateral Ground Clearance Check.......................................................................... 62 
Figure 48: Nose Gear Stick Diagram ......................................................................................... 63 
Figure 49: Main Gear Stick Diagram ......................................................................................... 63 
Figure 50: Nose (L) and Main (R) Gear ..................................................................................... 63 
Figure 51. Material Representation ........................................................................................... 64 
Figure 52: V-n Diagram............................................................................................................... 65 
Figure 53: Environmental impact Compared to Baseline ...................................................... 65 
Figure 54: Cabin temperature Control Zones .......................................................................... 66 
Figure 55: Mechanical Piping Schematic (Under Floor)......................................................... 67 
Figure 56: Air Distribution Schematic (Above Ceiling) ............................................................ 68 
Figure 57: Return Air & Exhaust Schematic............................................................................... 68 
Figure 58: Air Flow Visualization ................................................................................................. 68 
Figure 59: Environmental Control System Visualization........................................................... 69 
Figure 60: Noise Sensitivity to Flow Velocity.............................................................................. 70 
Figure 61: Sound Levels at Outlets ............................................................................................ 70 
Figure 62: Regenerative Noise Reduction Based On Plenum Design................................... 71 
Figure 63: Fuel System Schematic ............................................................................................. 73 
Figure 64: P-65 Wiring Layout ..................................................................................................... 74 
Figure 65: Pegasus 65 Flyaway Cost Curve.............................................................................. 76 
Figure 66: Flyaway Cost Breakdown By Individual Element................................................... 77 
Figure 67: Flyaway Cost Comparison to Similar Class Aircraft ............................................... 77 
Figure 68: Breakdown of O&M Costs ........................................................................................ 78 
Figure 69: Comparison of CASM Costs for Similar Class Aircraft............................................ 79

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Figure 71: Dimensioned 3 View ................................................................................................. 81 
List of Tables
Figure 1: Air Traffic RPK Growth Through 20245 .......................................................................... 2 
Figure 2: Air Traffic by Flow5.......................................................................................................... 3 
Figure 3: Regional Aircraft Proportion of World Fleets5 ............................................................. 3 
Figure 4: Average Gasoline Prices, 1991-201258 ........................................................................ 4 
Figure 5: Global LCC Growth11 .................................................................................................... 4 
Figure 6: Stakeholder Analysis...................................................................................................... 5 
Figure 7: Q40025 and CRJ70024..................................................................................................... 6 
Figure 9: Box Wing Configuration23 ........................................................................................... 10 
Figure 10: Turboprop21 and Propfan Engines28 ........................................................................ 12 
Figure 11: Propfan, Turboprop and Turbofan Propulsive Efficiency Plots31........................... 13 
Figure 12: Preliminary Design Configuration ............................................................................ 15 
Figure 13: Mission Profile ............................................................................................................. 20 
Figure 14: Weight Regression for Comparable Aircraft29 ....................................................... 23 
Figure 16: Comparison of Class I and Class II Drag Polar – Cruise......................................... 26 
Figure 17: Energy Based Constraint Sizing ................................................................................ 31 
Figure 18: Altitude vs. Fuel Burn for Multiple Mach at AR = 12 ............................................... 32 
Figure 19: Altitude vs. Fuel Burn for Multiple AR at M = 0.68 ................................................... 33 
Figure 20: Fuel Weight vs. Sweep Angle for Multiple AR......................................................... 33 
Figure 21: Final Wing Configuration .......................................................................................... 33 
Figure 22: Control Surface Diagram ......................................................................................... 37 
Figure 23: Longitudinal SAS Control39........................................................................................ 39 
Figure 24: SAS Response Properties39 ........................................................................................ 39 
Figure 25: CG Excursion Diagram.............................................................................................. 41 
Figure 26: Iterative Process Workflow........................................................................................ 42 
Figure 27: Climb Speed vs. Altitude .......................................................................................... 43 
Figure 28: Tail Forward on CG Travel for Multiple c (f/a)........................................................ 44 
Figure 29: Differential Alpha on CG for Multiple c (f/a) ......................................................... 44 
Figure 30: Sweep (f/a) on CG Travel for Multiple c (f/a)........................................................ 44 
Figure 31. P-65 Cabin Cross-Section ......................................................................................... 47 
Figure 32. Overall Fuselage Dimensions ................................................................................... 48 
Figure 33. Fuselage Interor Configuration ................................................................................ 48 
Figure 34. Fuselage Structure Model......................................................................................... 50 
Figure 35. Sheet-Stringer Approximation9................................................................................. 51 
Figure 36. Bay Structure (Dimensions in Inches)....................................................................... 52 
Figure 37. P-65 Cockpit Visibility................................................................................................. 53 
Figure 38. P-65 Cockpit17 ............................................................................................................ 53 
Figure 39: Blade Diameter vs. Blade loading........................................................................... 55 
Figure 40: Spin Rate vs. Power Loading per blade ................................................................. 56 
Figure 41: Final Engine Design ................................................................................................... 59 
Figure 42: Longitudinal Tip-Over Criterion Check47 ................................................................. 60 
Figure 43: Lateral Tip-Over Criterion47 ....................................................................................... 61 
Figure 44: Lateral Tip-Over Criterion Check ............................................................................. 61 
Figure 45: Longitudinal Ground Clearance Check................................................................. 61 
Figure 46: Lateral Ground Clearance Check.......................................................................... 62 
Figure 47: Nose Gear Stick Diagram ......................................................................................... 63 
Figure 48: Main Gear Stick Diagram ......................................................................................... 63 
Figure 49: Nose (L) and Main (R) Gear ..................................................................................... 63

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Figure 50. Material Representation ........................................................................................... 64 
Figure 51: V-n Diagram............................................................................................................... 65 
Figure 52: Environmental impact Compared to Baseline ...................................................... 65 
Figure 53: Cabin temperature Control Zones .......................................................................... 66 
Figure 54: Mechanical Piping Schematic (Under Floor)......................................................... 67 
Figure 55: Air Distribution Schematic (Above Ceiling) ............................................................ 68 
Figure 56: Return Air & Exhaust Schematic............................................................................... 68 
Figure 57: Air Flow Visualization ................................................................................................. 68 
Figure 58: Environmental Control System Visualization........................................................... 69 
Figure 59: Noise Sensitivity to Flow Velocity.............................................................................. 70 
Figure 60: Sound Levels at Outlets ............................................................................................ 70 
Figure 61: Regenerative Noise Reduction Based On Plenum Design................................... 71 
Figure 62: Fuel System Schematic ............................................................................................. 73 
Figure 63: P-65 Wiring Layout ..................................................................................................... 74 
Figure 64: Pegasus 65 Flyaway Cost Curve.............................................................................. 76 
Figure 65: Flyaway Cost Breakdown By Individual Element................................................... 77 
Figure 66: Flyaway Cost Comparison to Similar Class Aircraft ............................................... 77 
Figure 67: Breakdown of O&M Costs ........................................................................................ 78 
Figure 68: Comparison of CASM Costs for Similar Class Aircraft............................................ 79 
Figure 69: Advert for P-65 Marketing campaign ....................................................................... 1 
Nomenclature
a1 Regression Coefficient
A Regression Coefficient
AR Aspect Ratio
ax Acceleration in the X-direction
b Wing Span
B Regression Coefficient
b1 Regression Coefficient
c Chord
c1 Regression Coefficient
cd Two-Dimensional Coefficient of Drag
CD Three-Dimensional Coefficient of Drag
CD,0 Three-Dimensional Zero-Lift Coefficient of Drag
CD,flaps Three-Dimensional Coefficient of Drag for the Flaps
CD,fuselage Three-Dimensional Coefficient of Drag for the Fuselage
CD,nacelle Three-Dimensional Coefficient of Drag for the Nacelle
CD,tail Three-Dimensional Coefficient of Drag for the Tail
CD,wing Three-Dimensional Coefficient of Drag for the Wing
CG Center of Gravity
CGaft Most Aft Center of Gravity
CG-x Center of Gravity in the X-axis
cl Two-Dimensional Coefficient of Lift
Cl Coefficient of Rolling Moment
CL Three-Dimensional Coefficient of Lift
cl,mas Maximum Two-Dimensional Coefficient of Lift
CL,max Maximum Three-Dimensional Coefficient of Lift
Cm Coefficient of Pitching Moment
cmean Average Chord
Cn Coefficient of Yawing Moment

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cp specific fuel consumption
ctip Tip Chord
CY Coefficient of Side Force
d1 Regression Coefficient
D Drag
d/dt Derivative with Respect to Time
df Fuselage Diameter
e Oswald’s Efficiency Factor
E Endurance
ebpx Final Box Wing Oswald’s Efficiency Factor
etheo Theoretical Mono-Wing Oswald’s Efficiency
f Equivalent Parasite Area
f/a Forward/Aft
g Gravitational Acceleration
h Altitude
h/b Wing Separation as a Fraction of Wing Span
hcg Height of the Center of Gravity from the Ground
Iyy Area Moment of Inertia
K Drag Polar Coefficient of Lift Constant
ke,box Box Wing Correction Factor
ke,f Fuselage Effects Correction Factor
ke,M Mach Number Correction Factor
ke,WL Winglet Correction Factor
ke,Γ Dihedral Correction Factor
L Lift
L/D Lift to drag ratio
lf Fuselage Length
lm Distance between Aircraft Center of Gravity and Main Gear in the X-axis
ln Distance between Aircraft Center of Gravity and Nose Gear in the X-axis
M Mach
Mff Fuel Fraction
N Load Factor
np Number of Blades
NP Neutral Point
NPdes Desired Neutral Point
p Roll Rate
P/W Power to Weight Ratio
Pbl Power Loading per Blade
Pm Maximum Main Gear Strut Loading
Pmax Maximum Power
Pn Maximum Nose Wheel Strut Loading
Pn,dyn,t Maximum Dynamic Load Per Nose Gear Tire
Pnd Maximum Dynamic Nose Gear Load
PSL Power as Sea Level
q Pitch Rate
q∞ Dynamic Pressure
r Yaw Rate
R Range
Rfr Rolling Friction
S Wing Area
ST Stability Margin
STmin Minimum Stability Margin

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Sv Vertical Tail Area
Swet Wetted Wing Area
(t/c)r Thickness to Chord Ratio at the Root
V Velocity
VClimb Velocity of Climb
VS,L Stall Velocity for Landing
VS,TO Stall Velocity for Takeoff
Vtire,max Maximum Velocity of the Tire
Vv Volume Coefficient for Estimating Vertical Tail Size
W/S Wing Loading
WE Empty Weight
WF Fuel Weight
WF,res Reserve Fuel Weight
WL Landing Weight
WOE Operating Empty Weight
WPL Payload Weight
WTO Takeoff Weight
Wtfo Trapped Fuel and Oil Weight
Wcrew Crew Weight
xv Vertical Tail Sizing Characteristic Length
α Angle of Attack
αp Propulsive Efficiency
β Angle of Side Slip
βm Mass Fraction
Γ Dihedral Angle
η Efficiency
ηp propulsive efficiency
θ Longitudinal Ground Clearance Criterion
λ Taper Ratio
Λ Sweep Angle
λmin Minimum Taper Ratio
Λ(t/c)max Wing Sweep along Locus of Maximum Thickness
ρ∞ Freestream Air Density
σ Density Ratio
σallow Allowable Stress
τw Ratio of Thickness Ratio at the Tip to the Thickness Ratio at the Tip
φ Lateral Ground Clearance Criterion
ψ Lateral Tip-Over Criterion

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Introduction
Due to rising fuel costs, the plane ticket prices are also rising, which provides a barrier to
customers who have a desire to travel. This means that the market is ripe for a new regional
aircraft that can compete to an extent with the speed capabilities of the regional jet while
offering superior fuel savings. In their RFP, the AIAA has put out the requirements for a design that
responds to rising fuel prices2. The requirements of the RFP2 are as follows:
• Carry 75 passengers +/- 2.
• Fly an economic mission of 400 nmi at full capacity and a long range mission of
1600 at 90% capacity.
• Takeoff and land on 4,500 feet runways; takeoff and land on 8,000 foot runways
at 7,800 feet at temperature of 85°F at 80% Maximum Takeoff Weight (MTOW).
• Cruise at an altitude between 25,000 and 31,000 feet and a Mach number
between 0.62 and 0.68.
• Offer a fuel reduction of 65% from a similar class, existing regional jet.
• Offer a cost less than that of similar class, existing turboprops and significantly less
than that of similar class, existing regional jets.
• Jet-like experience and reduced noise levels compared to similar class, existing
turboprops.
• A turboprop engine design.
In the following sections, Pegasus Aerospace will develop its P-65 design.
Regional Jet Outlook
In order to understand the aircraft’s broad based applicability and the design decisions
that the Pegasus Aerospace team made, an understanding of the market and how the
stakeholders were considered is required.

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Market Analysis
Although short-term fluctuations exist in air travel, long-term estimates show a consistent
and robust growth in aviation driven by economic development, globalization and expanding
international trade, and liberalization. Airlines are constantly refining business models and the
internet provides added efficiencies for both passengers and airlines.
Economic growth is the principle driver of air transportation demand, and increases in
Gross Domestic Product (GDP) explain the majority of future growth. While the annualized
growth of GDP is expected to be 2.9%, air traffic is expected to exceed this, with annualized
growth rates of 4.8% and 6.2% for passenger and cargo flows respectively. These annualized
growth rates are visualized in Figure 1 and show a clear market need for new aircrafts. In fact, a
doubling of the world aircraft fleet is expected between 2005 and 2024, with 26,000 new aircraft
to be delivered at a value of $2.1 trillion (2004 USD).
FIGURE 1: AIR TRAFFIC RPK GROWTH THROUGH 20245
As shown in Figure 2, a considerable proportion of airline Revenue Passenger Kilometers
(RPKs) are generated by regional air travel, necessitating fleets of regional aircraft that
constitute a significant proportion of overall airline fleets.

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FIGURE 2: AIR TRAFFIC BY FLOW5
In 2004, regional aircraft constituted 15% of the world fleet, and as fleets grow, regional
aircraft will continue to be vital. In 2024, regional aircraft will account for a slightly larger 16% of
the world fleet, and as Figure 3 shows, the market will require new regional aircraft for fleet
expansion and replacement – approximately 2,900 aircraft. These market projections
demonstrate the clear need for a new regional aircraft in the near term.
FIGURE 3: REGIONAL AIRCRAFT PROPORTION OF WORLD FLEETS5

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The market need for regional aircraft tells only half the story. While the rising cost of jet
fuel, as seen in Figure 4, increases operational expenses of airlines, consumers within the industry
are demonstrating a growing air travel demand elasticity to airfare pricing. This is clear in the
growth of Low Cost Carriers (LCCs) as shown in Figure 5. A successful regional aircraft for the
near future will only satisfy market demand if it is demonstrably less reliant on the airlines’ fastest
growing cost – jet fuel.
FIGURE 4: AVERAGE GASOLINE PRICES, 1991-201258
FIGURE 5: GLOBAL LCC GROWTH11

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Stakeholder Analysis
A stakeholder mapping for the P-65 is shown in Figure 6 and shows a variety of
stakeholders and their power and interest. Primary stakeholders are seen to be passengers,
pilots, staff, shareholders, corporate leaders and employees, and airlines. Secondary
stakeholders include airports, competitors, near-airport residents, government, and energy
corporations.
FIGURE 6: STAKEHOLDER ANALYSIS
The valuation of stakeholders using the stakeholder analysis allowed for the
determination of the most necessary focus areas and characteristics in the creation of the P-65
during the preliminary design phase. The following are several key stakeholders determined using
this analysis: shareholders, corporate leaders, airlines, and passengers.
The importance of the shareholder, corporate leader, and airline lead to a focus on
bottom line expenses of the aircraft. Minimization of flyaway cost, cost per seat mile, and
operation and maintenance costs are key goals, as is manufacturability.

6
The importance of the consumer within the stakeholder analysis lead to a passenger-
centric engineering approach. Optimal fuel efficiency is a shared goal between all key
stakeholders, as reduced fuel use could result in reduced airfares and/or increased profit
margins. Pegasus Aerospace’s passenger-centric design mentality emphasizes an added
importance on the creation of a jet-like experience within our turboprop aircraft, leading to
some unconventional configuration choices.
Existing Designs
Some of the requirements of the RFP call for a reduction off of a baseline of existing
regional jet. Therefore it is important to understand what the baselines are. For existing regional
jets there are two main different types of regional jets: those with turboprop engines such as the
Bombardier Q400 and those with turbofan engines such as the Bombardier CRJ-700.
FIGURE 7: Q40025 AND CRJ70024
These two regional airlines are selected as baselines because they have roughly the
same passenger capacity as that of the RFP requirement and that is a key parameter for aircraft
class. The regional jet has generally fared better because it offers faster cruise speeds and
therefore reduced travel time when compared to the turboprop. In addition, customers often
stigmatize the turboprop because of its appearance and loud noise. But a turboprop does offer
an advantage in that it is more fuel efficient than the turbofan and therefore reduces fuel costs.
Both these relative advantages can be seen in Table 1.

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TABLE 1: Q4006 AND CRJ7003 SPECIFICATIONS
Aircraft Q400 CRJ700
Year Introduced 1984 2001
Takeoff Weight (lbs.) 64,500 72,750
Cruise Speed (kts./M) 360/0.54 447/0.78
Cruise Altitude (ft.) 25,000 37,000
Range (nm) 1360 1220
Passengers 68-86 66-78
Engine (x2) PW150A GE CF34-8C5B1
Maximum Fuel Weight (lbs) 11,700 19,600
Number Delivered 454 317
Configuration Selection
In order to select an overall configuration, a Figures of Merit (FOM) analysis is used to
evaluate different options.
Figures of Merit
The FOM analysis involves selecting parameters that are important to the design and
then assigning them a weight in the range of 1-5 (with 1 being the least important and 5 being
most important) based on how important each parameter is. Each aspect of the configuration is
then ranked on a scale of 1-5 (with 1 being the lowest and 5 being the highest) depending on
how well it performs in that category. Table 2 shows the identified FOM for the configuration
selection as well as their relative weights. As the call for 65% reduction in fuel consumption is
central to the RFP, it is given greatest priority. Drag reduction is mentioned as well because drag
reduction enables fuel consumption reduction. Innovation is also important because the public
has been reluctant to embrace the turboprop and a new design could help promote the
concept. Cost and airport compatibility are important because they are the more practical
drivers in the design and often stipulate whether or not airlines invest in the plane. Stall
characteristics are important for the practical design of the wing and whether or not the
configuration can sustain operation in its flight envelope. Passenger perception and
appearance are important, especially with the stigma against existing turboprops and the

8
requirement to simulate a jet-like experience. Finally capacity for fuel volume is a consideration
because some configurations offer more fuel storage than others.
TABLE 2: FIGURE OF MERITS CATEGORY WEIGHT
Figure of Merit Weight
Fuel/Drag Reduction 5
Innovation 4
Design Cost 4
Airport Compatibility 3
Stall Characteristics 3
Passenger Perception 1
Fuel Volume 1
The next section consists of a discussion of alternate configurations for three main
categories: fuselage and empennage design, wing design and engine system design. The merits
of these alternates are discussed and each section results in the choice of a design.
Fuselage/Empennage Alternatives
The fuselage and empennage are grouped together because certain configurations
under consideration are tailless, such as the blended wing body (BWB) and the flying wing. For
the conventional wing, different empennage options are offered.
Three main designs are considered for the fuselage/empennage group: conventional,
BWB and the flying wing, which are seen in Figure 8. The conventional design is the standard, the
design that all regional transports in existence have. It is the most well-known option and has the
greatest compatibility with manufacturing capability and airports, thus giving it some inherent
cost benefits. Not much is said in the proceeding section about the conventional wing; rather it
is used as a standard with which to compare the BWB and flying wing designs.

9
As compared to the conventional
design, both the flying wing and the BWB
offer improved aerodynamics, mainly in
reduced drag47. This translates to a lower
necessary power-to-weight ratio (P/W)
which leads to greater fuel efficiency and
a lower weight. However, the flying wing
and the BWB have stability issues due to
their lack of a tail, though this can be
combatted with modern flight controls.
Both the flying wing and the BWB are
innovative but their cabin layouts
detrimentally affect the experience of the
passengers because not every passenger
can have a window seat and the cabin is
potentially cavernous. As the BWB and
flying wing designs are important to the
discussion on wing selection, they are
discussed in more detail in that section.
As far as empennage selection
accompanying the conventional
fuselage, two options are considered: the
conventional tail and the T-tail. For the conventional tail, the location of the horizontal tail in
relation to the vertical tail is all the way on the bottom, giving the conventional tail an
advantage in maintenance cost when compared to the T-tail. The primary advantage of the T-
tail is that the high location of the horizontal tail means less interference caused by the wing is
FIGURE 8: CONVENTIONAL (TOP)24, BWB
(MID)21 AND FLYING WING DESIGNS (BOT)27

10
experienced by the horizontal tail, which lead to higher efficiencies47. Additionally, a T-tail puts
more of a structural burden on the vertical tail, requiring it to be both stronger and heavier.
Because of the heavy coupling between the fuselage and wing selection, this section
does not result in an evaluation of the FOM matrix. Rather, this is deferred to the section on the
wing.
Wing Alternatives
For the wing selection, in addition to the BWB and the flying wing, a conventional wing
and a box wing are considered.
FIGURE 9: BOX WING CONFIGURATION23
From the perspective of innovation, the conventional wing design offers the least
innovative option for the aircraft. While the conventional wing is extremely reliable and
effective, newer design configurations can provide performance improvements. Flying wings
offer a slightly more innovative design; however, examples have already been developed and
flown in the past. The BWB and box wing configurations are the most novel designs among those
considered. The box wing design provides a slight innovation advantage over the blended wing
due to the lack of test data for box wing aircraft. With increased innovation comes increase in
cost necessary to develop emerging technologies to requisite safety levels. Therefore, the cost
FOM scores are inversely proportional to innovation scores.

11
Conventional wings offer the most information and data for stall characteristics due to
their widespread application in aviation, which gives it an advantage over the others. Despite
the lack of actual data, box wing designs can be tailored to provide favorable stall
characteristics. The forward wing can be designed such that it stalls before the rear wing,
resulting in a nose down moment that leads to recovery. Blended wing and flying wing designs
have the least favorable stall characteristics, but the blended wing configuration offers a slight
stall characteristic advantage over flying wings, which results from the aerodynamic effects and
weight distribution caused by the protruding fuselage structure.
Selecting a wing design that is compatible with the majority of commercial airports
eliminates the need to extensively modify the terminal and gate configuration. Conventional
wings score the highest in this category because current airports were designed around
conventional winged aircraft. Box wings would not significantly interfere with airport operations
as they have approximately the same compatibility as a conventional wing. Blended wing
designs result in more accessible surfaces for doors, but would require some modification to the
gate orientation. A flying wing design complicates the loading and unloading of passengers
and luggage due to its shape, and is the least compatible.
The configuration with the highest inherent lift-to-drag ratio (L/D) is the flying wing. The
inherent L/D values for the blended wing-body and the box wing are fairly similar and slightly
lower than the flying wing. All three of these have greater drag reduction than the conventional
wing.
Internal fuel volume is relatively negligible in the selection of the wing configuration as
the wing can be scaled to fit as much fuel as required; however, it can be difficult to create
sufficient wing fuel capacity in the BWB and flying wing configurations, so a consideration is
necessary. Box wings and conventional wings have comparable fuel volume, as the box wing
has a comparable wing area to the conventional wing, with the area spread over two wings
rather than one.

12
Table 3 shows the FOM analysis for the fuselage/wing category, showing that the flying
wing is the selected design moving forward.
TABLE 3: WING/FUSELAGE FOM DETERMINATION
Figure of Merit Conventional Box Wing BWB Flying Wing
Fuel/Drag Reduction 2 4 4 5
Innovation 1 5 4 3
Design Cost 5 1 2 3
Airport Compatibility 5 5 3 1
Stall Characteristics 5 3 2 1
Passenger Perception 3 5 5 5
Fuel Volume 4 4 2 2
Sum 71 77 66 62
Engine Parameter Alternatives
In selecting the engine type, the two choices are the traditional turboprop and the
propfan, shown in Figure 10.
FIGURE 10: TURBOPROP21 AND PROPFAN ENGINES28
Both of these engines fall under the category of “turboprops” and thus fall within the
requirements for propulsion as stated in the RFP. The main difference between the propfan and
the turboprop is that the blades of the propfan are shorter and greater in number than the
blades of the turboprop. This difference allows the propfan to avoid transonic drag divergence
problems at the tips of its blades when flying at higher Mach numbers. As seen in Figure 11, the
envelope where propfans achieve their highest efficiency is between Mach 0.6 and Mach 0.7.
Traditional turboprops are far less efficient in these flight regimes. The RFP requires that the
cruising Mach number of the aircraft lie between Mach 0.62 and 0.68, which is in the envelope
of maximum efficiency for the propfan.

13
FIGURE 11: PROPFAN, TURBOPROP AND TURBOFAN PROPULSIVE EFFICIENCY PLOTS31
Since the turboprop is more widespread, it is easier to maintain and operate, given the
data and industry knowledge that existed. Propfans also incur a weight penalty relative to the
turboprop due to their numerous fan blades and additional gearing. Finally, the turboprop is
considered to be more attractive to the consumer due to the fact that it is more traditional and
the propfan’s blades appear dangerous. However, despite the advantages of the turboprop, as
seen in Error! Not a valid bookmark self-reference., the propfan is selected because it fits
well with the cruise conditions as outlined in the RFP.
TABLE 4: ENGINE TYPE FOM DETERMINATION
Figure of Merit Turboprop Propfan
Fuel/Drag Reduction 2 5
Innovation 3 5
Design Cost 3 3
Airport Compatibility 5 5
Stall Characteristics 3 3
Passenger Perception 4 3
Fuel Volume 3 3
Sum 65 87
Next, the number of engines is determined. Only one or two engines is considered as
three or more engines are considered to be excessive for the mission requirement. Table 5 shows

14
that two engines are favored over one because the safety in redundancy they add outweighs
the weight and cost savings that one engine would provide.
TABLE 5: NUMBER OF ENGINES FOM DETERMINATION
Figure of Merit One Engine Two Engines
Innovation 3 3
Design Cost 4 3
Fuel Volume 3 3
Sum 72 75
Another consideration for the engine selection is that of placement. The choice of
engine placement is between a forward location with engines mounted on the forward wing
and an aft location with the engines mounted on the rear of the fuselage. Based on the RFP,
noise is an important issue, particularly as it relates to passenger perception and that proves to
be the deciding factor for the aft located engines, as can be seen in Table 6. Due to the
importance of passenger perception, it is logical to push back the location of the engines are
far back as possible so as to offer maximum separation between the passengers and the
engines.
TABLE 6: LOCATION OF ENGINES FOM DETERMINATION
Figure of Merit Forward Aft
Innovation 3 3
Design Cost 3 3
Fuel Volume 3 3
Sum 67 70
The final engine selection choice is deciding between the pusher and puller
configuration. The main advantage of the pusher configuration is its potential to dramatically
reduce cabin noise. Keeping the engines’ blades further back also is beneficial in terms of

15
safety. As a more conventional configuration, the puller configuration enjoys a slight advantage
in cost and perception. Table 7 shows that ultimately the pusher configuration is selected.
TABLE 7: CONFIGURATION OF ENGINES FOM DETERMINATION
Figure of Merit Pusher Puller
Innovation 4 3
Design Cost 3 4
Fuel Volume 3 3
Sum 75 74
Configuration Selection
In summary, the configuration of the P-65 consists of a box wing with two propfan
engines mounted on the rear of the fuselage. Figure 12 gives a visualization of the preliminary
design.
FIGURE 12: PRELIMINARY DESIGN CONFIGURATION

16
Technologies
In order to reach the required 65% fuel consumption, a number of technologies are
needed. Several existing and emerging technologies are listed below and are compared for
compatibility. A final technology matrix is then assembled.
Box Wing
A box wing is a configuration comprised of two sets of horizontal wings that are
connected by vertical tips. It has an advantage in comparison to a conventional cantilevered
design in that it has higher stiffness, lower induced drag, a higher trimmed maximum three
dimensional (3D) coefficient of lift (CLmax) and reduced wetted area and therefore less parasite
drag. All of these things contribute to a reduced weight compared to the conventional
configuration. Wolkovitch56 estimates that the box wing reduces the equivalent
FIGURE 13: CONFIGURATION 3 VIEW

17
wing/empennage weight 20 to 40%, depending on the aspect ratio, sweep and other factors. In
addition, the span efficiency factor is improved 40 to 75% through the box wing.
Propfan
A propfan is a jet engine that has elements of both a turboprop and a turbofan. It is like
a turboprop in that it has a propeller; however the function of the propeller is similar to that of a
turbofan bypass compressor. Alternate names for the propfan the ultra-high bypass turbofan.
The design of the turbofan is intended to combine the advantages of the turboprop - fuel
economy - with the advantages of the turbofan - performance. The propfan reduces
compressibility losses, thus increasing fuel efficiency. Research done by Rolls-Royce53 suggests an
improvement in Specific Fuel Consumption (SFC) of 30% for the propfan over the turbofan.
Composite Structures
Composite structures have an advantage compared to traditional metal structures in
that they have a lower density and a higher specific strength, which combine to reduce
structural weight. Composites are more customizable in terms of strength and stiffness. However
there are some costs to take into account. As composites are relatively new, the manufacturing
capability is not as advanced as it is for more traditional materials. The material itself is expensive
and can be hard to implement. Weighing these pros and cons, the proposed design has a
composite fuselage and aluminum wings. For a fuselage, Roskam50 estimates that switching from
aluminum to composites results in a weight savings of 15 to 25%.
Natural Laminar Flow
NLF is an attempt to prolong the laminar flow of air over different parts of the aircraft
through aerodynamic shaping. This is beneficial as laminar flow has higher coefficient of lift
capability than turbulent flow, as well as reduced skin friction drag. The P-65 incorporates NLF in
the nose and wing. The nose must be carefully shaped and the wing maintains laminar flow
through the careful selection of an airfoil. The airfoil pushes the transition point from laminar to
turbulent further back on the airfoil and therefore maintains the favorable laminar region as long

18
as possible. NLF offers a potential 25% drag reduction, if fully integrated in the entire aircraft,
according to National Aeronautics and Space Administration (NASA) research19.
Digital fly-by-wire
Using a digital fly-by-wire control offers weight savings when compared to hydraulic
controls because they are lighter and take up less volume. The International Air Transport
Association (IATA)20 estimates a 1 to 3% fuel savings for the use of digital fly-by-wire.
Solid Oxide Fuel Cell Auxiliary Power Unit
Solid Oxide Fuel Cell (SOFC) Auxiliary Power Units (APU) generate electrical power more
efficiently than their existing counterparts. They potentially offer a fuel savings of 40% during
startup13.
Riblets
Riblets are grooves running lengthwise on the surface of aircraft that reduce skin-friction
drag for turbulent airflow. Historically, riblets had a drawback in that the film they were molded
into only lasted a few years. Now a new riblet technique has been developed that results in
greater longevity without adding to the aircraft weight. The University of Illinois Aerospace
Department52 estimates a reduction in drag of 1 to 4% for the use of riblets.
Spyroid Wingtip
The spyroid wingtip is a wingtip with a closed box shape. It offers the benefits of a box
wing to a reduced effect, particularly in terms of reduced drag. The technology is estimated to
have a drag reduction of 11%38.
Circulation Control Wing
A circulation control wing (CCW) is a high-lift device that increases the lift of the wing
through increasing the velocity of the flow over the leading and trailing edges34. It replaces
traditional flaps and slats, which introduce extra drag with the additional lift provided. CCW
reduces wing structure weight and allows for better takeoff and landing performance. This leads

19
to a reduction in fuel needed during these sections. However the CCW is a complicated
technology that requires power from the engine which reduces its efficiency.
Technology Compatibility
As a final step the technologies are evaluated as a whole to determine if they are
compatible. For the most part the different evaluated technologies do not overlap in their effect
on the aircraft so they are compatible by default. For obvious reasons, the box wing and the
spyroid wingtips cannot both be used. CCW is a complicated technology that causes difficulty
both with the wing and with the efficiencies of the propfan. Table 9 shows the technology
compatibility matrix that is used to decide on a final list of technologies that the P-65 will use.
TABLE 9: SELECTED
TECHNOLOGIES
P-65 Technologies
Box Wing
Propfan
Composites
NLF
Riblets
Digital Fly-by-wire
SOFC APU
TABLE 9: P-65 TECHNOLOGY COMPATIBILITY MATRIX
O = incompatible
X = compatible
BoxWing
Propfan
Composites
NLF
DigitalFly-by-wire
SOFCAPU
Riblets
SpyroidWingtip
CCW
Box Wing X X X X X X O O
Propfan X X X X X X O
Composites X X X X X X
NLF X X X X X
Digital Fly-by-wire X X X X
SOFC APU X X X
Riblets X X
Spyroid Wingtip X
CCW

20
Overall, the box wing and NLF technologies are valued over their alternatives, which
leads to the finalized P-65 technology package listed in Table 9.
Weight Sizing
The aircraft’s mission, performance objectives, and payload dictate the initial sizing
process so that it meets all of the necessary requirements. The following weight sizing for the P-65
is done using the methods provide in Airplane Design: Part 144 and provides initial estimates for
the empty weight (WE) and gross takeoff weight (WTO).
Mission Profile
The RFP provides a sizing mission with a 1600 nmi range at a 90% load factor, along with a
typical economic mission of 400 nmi with a full load of 74 passengers. The weight sizing process
uses the longer design mission for flight phase calculations. Figure 14 illustrates the two missions
and indicates the different flight phases.
FIGURE 14: MISSION PROFILE
Weight Estimation
Weight Convergence Method
The P-65’s primary role is regional passenger transportation. Tailoring the performance for
cruise enables the vehicle to achieve a lower fuel burn over the longest phase in the mission
profile. The takeoff weight is comprised of the mission fuel weight (WF), payload weight (WPL)
and the operating empty weight (WOE). The operating empty weight itself is comprised of the
empty weight, the trapped fuel and oil weight (Wtfo) and the crew weight (Wcrew).

21
The fuel weight results from fuel-fraction calculations and an initial gross takeoff weight
estimate. Subtracting out the payload, trapped fuel and oil, and fuel weights from the initial
estimate results in the empty weight of the aircraft. An iterative process varies the takeoff weight
estimateuntil the calculated empty weight and the allowed empty weight determined from
takeoff weight regression converge. This process only works if the L/D values calculated using the
drag polar approximation are reasonable estimates.
Payload Weight
The payload associated with commercial airline transportation includes passengers and
luggage. As specified in the RFP, a single passenger with luggage weighs 225 lbs. The P-65 has
74 seats, resulting in a maximum payload of 16,650 lbs.
Mission Fuel Weight
The mission fuel weight accounts for all of the fuel used for the design mission profile. The
fuel-fraction method represents the fuel used during each phase as a ratio of the weight of the
aircraft at the end of the phase to the weight of the aircraft at the beginning of the phase.
Equation 1 shows the fuel-fraction (Mff) derivation used for the climb and descent phases using
Breguet’s endurance equation. Equation 2 shows a similar calculation for the cruise phase using
Breguet’s range equation. Lift-to-drag ratio (L/D) values are obtained from the class II drag polar
and the engine deck data provides the appropriate SFC (cp) and propulsive efficiency (ηp). The
velocity (V) is also an important consideration for the two equations.
𝐸 =
375(𝐿/𝐷)𝜂!
𝑉𝑐!
ln (
1
𝑀!!
)
EQUATION 1: BREGUET’S ENDURANCE EQUATION
𝑅 =
375(𝐿/𝐷)𝜂!
𝑐!
ln (
1
𝑀!!
)
EQUATION 2: BREGUET’S RANGE EQUATION
Table 10 shows the calculated fuel fractions, start weights, and the fuel used for each
phase.

22
TABLE 10: MISSION PHASE WEIGHTS AND FUEL FRACTIONS
Segment Mff Start Weight (lbf) Fuel Burn (lbf)
Start/Taxi 0.9980 45250.3 90.5
Take-off 0.9950 45159.8 225.8
Climb 0.9896 44934.0 466.7
Cruise 0.9221 44467.3 3464.4
Descent 0.9967 41002.9 133.8
Land 0.9950 40869.1 204.3
Reserve 0.9800 40664.8 813.3
Total 0.8807 39851.5 5398.8
The fuel fraction approximations for the start/taxi, take-off, landing, and reserve phases
result from historical data for regional turboprop aircraft. The total mission fuel-fraction results
from multiplying the fuel-fractions for each segment. The total fuel weight is then determined by
Equation 3, which accounts for the weight of any fuel reserves (WF,res).
𝑊! = 1 − 𝑀!! 𝑊!" + 𝑊!,!"#
EQUATION 3: FUEL WEIGHT
Gross Takeoff and Empty Weights
Historical data for aircraft of comparable missions and configurations provides a way to
estimate the gross takeoff weight and empty weight of the new P-65 aircraft. Equation 4 gives
the regression relationship between the takeoff and empty weight. Linear regression of the data
provides a means to calculate the slope and y-intercept of the fitted curve, which allows for
weight interpolation. The calculated mission fuel fractions, drag polar L/D’s, and payload
weights allow for the calculation of the aircraft empty weight. The estimated takeoff weight is
varied while comparing the calculated empty weight to that obtained from the weight
regression.
log!" 𝑊! =
log!" 𝑊!" − 𝐴
𝐵
EQUATION 4: TAKEOFF AND EMPTY WEIGHT REGRESSION
Figure 15 shows the weight regression analysis performed on 42 subsonic turboprop and
regional jet transport aircraft, where the constants A and B can be determined from the

23
regression line. The values for A and B, as well as the method for calculating them, can be found
in Table 6.
FIGURE 15: WEIGHT REGRESSION FOR COMPARABLE AIRCRAFT29
TABLE 11: CONSTANT VALUES - WTO VERSUS WE
Coefficient Formula Value
A y-int*(-1/slope) 0.14678
B 1/slope 1.018837
Drag Polar (Class I Method)
The Class I method is used for determining the drag polar at low speeds and is taken from
Airplane Design: Part I44. Equation 5 gives the drag coefficient (CD) of an aircraft based on the
parabolic assumption of the drag polar where the non-zero lift drag coefficient (CD,0) can be
expressed as a function of equivalent parasite area (f) and the wing area (S). The constant K is a
function of the aspect ratio (AR) and Oswald’s efficiency factor (e).
𝐶! = 𝐶!,! + 𝐾𝐶!
!
, 𝑤ℎ𝑒𝑟𝑒 𝐶!,! =
𝑓
𝑆
, 𝐾 =
1
𝜋𝑒𝐴𝑅
EQUATION 5: CLASS I DRAG POLAR
In order to determine f, Equation 6 gives the relationship between equivalent parasite
area and wetted wing area (Swet). In a similar manner, Equation 7 gives the relationship between
wetted wing area and takeoff weight relates. The constants a1, b1, c1 and d1 are taken from
historical data and are given in Table 11.

24
log!" 𝑓 = 𝑎! + 𝑏!log!" 𝑆!"#
EQUATION 6: EQUIVALENT PARASITE AREA
log!" 𝑆!"# = 𝑐! + 𝑑!log!" 𝑊!"
EQUATION 7: WETTED WING AREA
TABLE 12: EQUIVALENT PARASITE AND WETTED WING AREA REGRESSION
COEFFICIENTS
Parameter Value
a -2.5229
b 1.0
c -0.0866
d .08099
Table 13 gives a list of parameters used in the Class I drag polar, culminating in values for
K and CD,0.
TABLE 13: CLASS I DRAG POLAR PARAMETERS
Parameter Value
W/S (lb/ft2) 81
AR (-) 12
e (-) 1.44
Swet (ft2) 5197.7
WTO (lbs) 45250
f (ft2) 15.6
K (-) 0.0184
CD,0 (-) 0.0279
Average velocity, altitude, and phase weight, as well as wing area, provide an average 3D
coefficient of lift (CL) for the climb, cruise, and reserve phases. Using the CL values and data
from Table 13, Equation 5 yields values for CD. Equation 8 shows how the CL and CD relate to the
L/D. Table 14 gives the L/D for each phase of interest.
𝐶!
𝐶!
=
𝐿
1
2 𝜌! 𝑉! 𝑆
𝐷
1
2
𝜌! 𝑉! 𝑆
=
𝐿
𝐷
EQUATION 8: L/D

25
TABLE 14: PHASE SPECIFIC L/D
Phase L/D (-)
Climb 20.6
Cruise 12.8
Decent 13.2
Drag Polar (Class II Method)
The class I drag polar determination is used in the initial iteration of design and sizing of
the P65, while a class II method is utilized for more refined drag estimates. The class II drag
estimates are used in design iterations until vortex lattice provide progressively more accurate
drag estimations.
Class II Parasitic Drag
The class II parasitic drag considers friction, profile,
interference and excrescence drag through component-
based empirical relationships. A decomposition of the parasitic
drag based on the aircraft’s components is given in
Equation 9. Decomposition is configuration dependent
and contributions of each component are scaled based
upon reference area.
𝐶! = 𝐶!,!"#$ + 𝐶!,!"#$%&'$ + 𝐶!,!"#$ + 𝐶!,!"#$%%$ + 𝐶!,!"#$%
EQUATION 9: DRAG DECOMPOSITION INTO
COMPONENTS
The zero lift drag coefficients for these
components (except for flaps) are found using
Roskam’s method for estimating drag polar51 and are
given in Table 15. Figure 16 shows the breakdown for
each components contribution to zero lift drag.
FIGURE 16: ZERO LIFT DRAG
COEFFICIENTS BREAKDOWN
TABLE 15: ZERO LIFT DRAG
COEFFICIENTS
Component CD,0
Fuselage .0110
Nacelles .000241
Wings .00875
Tail .00622
Total .0262

26
The unique wing, nacelle, and tail Reynolds numbers are all varied with altitude so that
the drag polar for each mission segment is as accurate as possible.
Wave drag is assumed to be negligible at the P-65’s design cruise speed and induced
drag is approximated using Oswald’s efficiency factors validated in literature.
Figure 17 offers a comparison between the class I and class II drag polar methods.
FIGURE 17: COMPARISON OF CLASS I AND CLASS II DRAG POLAR – CRUISE
Constraint Sizing
The goal of constraint sizing is to determine the power to weight ratio and the wing
loading of the aircraft. These are attained by analyzing point energy constraint equations; the
most basic of which is given in Equation 10. This equation is a function of power at sea level (PSL),
the maximum takeoff weight, the current mass fraction (βm) of the aircraft, the propulsive
efficiency (αp), the freestream velocity, the freestream dynamic pressure (q∞), the wing area, the
zero lift coefficient of drag, K, the load factor (N), the rolling fiction (Rf), the altitude (h) and the
gravitational acceleration (g) for earth.
𝑃!"
𝑊!"
=
𝛽!
𝛼!
𝑉
𝑞 𝑆
𝛽 𝑊!"
𝐶!,! + 𝐾
𝑁 𝛽! 𝑊!"
𝑞! 𝑆
!
+
𝑅!
𝑞! 𝑆
+
1
𝑉
𝑑
𝑑𝑡
ℎ +
𝑉!
2𝑔
EQUATION 10: ENERGY CONSTRAINT55

27
The inputs for this equation are distinct for each set of flight conditions; however, the
values which determine some of these inputs are constant across all flight conditions of the
same configuration. The flight conditions of interest for the constrain sizing are: climb cruise,
operational ceiling and high altitude and sea level takeoff and approach.
K Determination
The sweep value determines the optimal taper ratio to attain an elliptical lift distribution
according to Equation 12, which is a function of sweep angle and taper ratio (λ). This makes use
of the minimum taper ratio (λmin) found in Equation 11. Equation 12 is a piecewise function as a
tip chord of 1.5 feet is deemed the minimum permitted tip chord for this box wing because of
the perceptual constraints from the market analysis.
𝜆!"# = −
𝑐!"#
𝑐!"# − 𝑐!"#$
EQUATION 11: MINIMUM TAPER RATIO GIVEN TIP CHORD36
𝜆 =
𝜆!"#, 0.45𝑒!!.!"#$∗!
< 𝜆!"#
0.45𝑒!!.!"#$∗!
, 0.45𝑒!!.!"#$∗!
≥ 𝜆!"#
EQUATION 12: TAPER RATIO36
That taper ratio is used in conjunction with the aspect ratio to determine the theoretical
mono-wing Oswald efficiency (etheo) as shown in Equation 13.
𝑒!!!" =
1
(1 + 𝐴𝑅(0.0524𝐴𝑅! − 0.15𝐴𝑅! + 0.1659𝐴𝑅! − 0.0706𝐴𝑅 + 0.0119))
EQUATION 13: MONO-WING THEORETICAL OSWALD FACTOR36
There is an inherent loss of efficiencies due to interference effects, which is covered by
another efficiency term, ke,Do. This value is given as 0.86436.
The fuselage area also plays a part in the total Oswald efficiency, where the fraction of
the span which is occupied by the fuselage necessitates another correction term, ke,f, as
defined in Equation 14.
𝑘!,! = 1 − 2
𝑑!
𝑏
!

28
EQUATION 14: FUSELAGE EFFECTS CORRECTION FACTOR36
The next correction is for dihedral (Γ). Dihedral does not change the reference area of
the wing; however, wing with dihedral is longer than the zero-dihedral wing and therefore attains
slightly higher efficiency levels than a wing with no dihedral. This correction factor, ke,Γ, is given in
Equation 15, where the dihedral angle is for each wing.
𝑘!,! =
1
cos(2𝛤)!
EQUATION 15: DIHEDRAL CORRECTION FACTOR36
Winglets also necessitate a correction term. Winglets increase the performance of
aircraft and this is captured in terms of another correction factor, ke,WL, as given by Equation 16,
where h/b is the wing separation as a fraction of the wing span.
𝑘!,!" = 1 +
2
2.83
ℎ
𝑏
!
EQUATION 16: WINGLET CORRECTION FACTOR36
Mach effects also need a correction. As Mach number increases above a value
of 0.3, drag increases at a significantly higher rate than it does in the incompressible regime. The
correction factor, ke,M, takes this non-linearity into account and is given in Equation 17.
𝑘!,! =
1, 𝑀 < 0.3
−0.001521 ∗
𝑀
0.3
− 1
!".!"
+ 1, 𝑀 ≥ 0.3

EQUATION 17: MACH NUMBER CORRECTION FACTOR36
There is one remaining correction factor. This factor, ke,box, takes the interaction of the
two wings into account. This variable depends on several constants and the mean separation
between the wings. The results of this factor give the box wing its particular edge over
conventional designs.
𝑘!,!"# =
𝑘3 + (𝑘4 ℎ
𝑏)
𝑘1 + (𝑘2 ℎ
𝑏)
; 𝑤ℎ𝑒𝑟𝑒
𝑘1
𝑘2
𝑘3
𝑘4
=
1.037
0.571
1.037
2.126
EQUATION 18: BOX WING CORRECTION FACTOR36

29
Once all of the correction factors are computed, they are all multiplied together to
obtain the final box wing Oswald efficiency factor (ebox) for the aircraft, as seen in Equation 19.
This factor is combined with the aspect ratio to yield the first order drag coefficient, as shown
below.
𝑒!"# = 𝑒!!!" 𝑘!,!! 𝑘!,! 𝑘!,! 𝑘!,!" 𝑘!,! 𝑘!,!"#
EQUATION 19: FINAL OSWALD EFFICIENCY FACTOR36
𝐾 =
1
𝜋 𝑒!"# 𝐴𝑅
EQUATION 20: FIRST ORDER DRAG COEFFICIENT
Constraint Sizing Results
The input terms for the energy constraint equation are shown in Table 16 and Table 17.
The energy constraint equation must be manipulated to attain takeoff and landing parameters;
however, as it is a simple algebraic manipulation, it is not shown here.
TABLE 16: ENERGY EQUATION INPUTS 1
Variable β (-) σ (-) αp (-) q
(lb/ft3)
Runway
Length (ft)
Safety
Factor
Derivation Wsegment/WTO ρ/ρSL σ0.7 ½ρV2 Defined Defined
MissionSegment
Takeoff 0.998 1.000 1.000 ~ 4000 1.1
High Altitude
Takeoff
0.798 0.689 0.770 ~ 8000 1.1
Climb 0.993 0.361 0.490 94.81 ~ ~
Cruise 0.983 0.361 0.490 194.71 ~ ~
Operational
Ceiling
0.983 0.361 0.490 94.81 ~ ~
Approach 0.998 1.000 1.000 ~ 4000 1.3
High Altitude
Approach
0.798 0.689 0.770 ~ 8000 1.3

30
TABLE 17: ENERGY EQUATION INPUTS 2
Variable S (ft2) N (-) R (-) dh/dt
(ft/sec)
dV/dt
(ft/sec2)
K1 (-)
Derivation Variable Constant Constant Defined Defined See Above
MissionSegment
Takeoff ~ 1 0 0 0 0.0184165
High Altitude
Takeoff
~ 1 0 0 0 0.0184165
Climb ~ ~1 0 15.04 0 0.0184166
Cruise ~ 1 0 0 0 0.0187853
Operational
Ceiling
~ ~1 0 15.04 0 0.0184166
Approach ~ 1 0 0 0 0.0184165
High Altitude
Approach
~ 1 0 0 0 0.0184165
When the above values are applied and the results are plotted against various wing
loadings, Figure 18 is the result of the constrain equation applied to each condition of interest.
The viable region of this plot is to the left of both vertical lines and above the highest horizontal
line. As there may be some modifications to the design necessary at a later time or allowing for
the possibility that the operator may decide to operate the aircraft outside the envelope given
in the RFP, slight excess power is required. Furthermore, due to the possibility of gusts while
landing, it is inadvisable to operate an aircraft too close to stall. Due to these reasons, the point
shown in Figure 18 is selected as the operating point, which is specified in Table 18.

31
FIGURE 18: ENERGY BASED CONSTRAINT SIZING
TABLE 18: ENERGY BASED CONSTRAIN SIZING RESULTS
P/W W/S (lb/ft2)
0.18 81
As stated previously in the weight sizing section, this process was performed iteratively
while mapping the design space. As a result, the plot shown in Figure 18 is for the chosen
configuration and is distinct for most cases within the design space.
Wing Design
Planform Selection
The planform includes aspect ratio, wing loading, sweep, and taper ratio. The wing
loading is determined by the constraint sizing section and is 81 lb/ft2. The sweep angle is
determined by the neutral point considerations, which are discussed in the stability and control

32
section. The sweep angle determines the taper ratio, as seen in the constraint sizing section.
Thus, the planform depends primarily on aspect ratio and sweep angle. As can be seen in Figure
19 and Figure 20, the relative decrease in fuel consumption associated with increasing AR
decreases as AR grows. Due to this fact, as well as considering structural effects, an AR of 12 is
selected.
FIGURE 19: ALTITUDE VS. FUEL BURN FOR MULTIPLE MACH AT AR = 12

33
FIGURE 20: ALTITUDE VS. FUEL BURN FOR MULTIPLE AR AT M = 0.68
The effects of sweep angle on fuel use are not as easy to see. However, it is evident from
Figure 21 that the lower the amount of sweep, the lower the fuel usage. Ultimately, sweep angle
is primarily driven by stability concerns.
FIGURE 21: FUEL WEIGHT VS. SWEEP ANGLE FOR MULTIPLE AR
The final wing configuration is shown in Figure 22.
FIGURE 22: FINAL WING CONFIGURATION

34
Airfoil Selection
As discussed in the technology section, the aircraft makes use of NLF airfoils. NACA 6-
and 7-series airfoils incorporate NLF because the thickest part of the airfoil is significantly farther
aft than on the 4-series. As the basic constraint sizing had already been performed, the cruise
two-dimensional (2D) coefficient of lift (cl) is known to be approximately 0.4. Thus, the airfoils
under consideration must maintain the drag bucket through at least this point, if not beyond.
Furthermore, cl values for takeoff and landing are 2.1 and 2.5, respectively. As a result, the airfoil
not only has to maintain the drag bucket through a cl of 0.4, it also must attain a maximum cl
(cl,max) with a pivot flap of at least 1.4 (obtained by comparing different flaps for 2 different
airfoils, NACA 63(4)-420 and NACA 65(3)-118). This results in 5 possible airfoils for wing use. Their
properties are shown in Table 19, and the NACA 65-4010 is the selected airfoil.
TABLE 19: AIRFOIL SELECTION OPTIONS1
Airfoil cl,max cd@cl=0.4 Min
Moment
Max
Moment
Δ
Moment
CL,max
(20
deg)
CL,max
(45
deg)
63(1)-412 1.75 0.0055 -0.6 -1.1 0.5 2.5 3.38
64(1)-412 1.6 0.0048 -0.85 -1.1 0.25 2.35 3.23
65-410 1.5 0.0045 -0.9 -1.1 0.2 2.25 3.13
65(1)-412 1.6 0.0047 -0.9 -1.15 0.25 2.35 3.23
66(2)-415 1.6 0.0045 -0.8 -1.35 0.55 2.35 3.23
The NACA 65-410 was chosen above the other airfoils due to its superior 2D coefficient of
drag (cd) and low change in moment across its useful angle of attack envelope. The airfoil
selected for the winglets and tail is the NACA 65-009 due to its low drag over a small bucket
while maintaining laminar flow over a relatively large portion of the airfoil.
Empennage Sizing
As the box wing does not have a horizontal tail attached to the empennage, only the
design of the vertical tail is considered. The sizing process for this component is performed using
Airplane Design: Part II47. These equations rely on the distance (xv) between the center of gravity

35
(CG) and the aerodynamic center of the stabilizing surface, the relative area of the stabilizing
surface (Sv) to the main wing, and the wing span to determine the volumetric constant. These
constants have been calculated for many aircraft and can be used as a guideline for
empennage sizing before stability analysis takes place; the average for turboprops and regional
airliners yields a resultant Vv of 0.0847. Furthermore, the relatively large winglets contribute to the
stabilizing area, as does the 7.0 degree dihedral of each wing, which requires less total area
being for the primary vertical tail. Equation 21 is used to determine the tail characteristic sizes.
𝑉! =
𝑥! 𝑆!
𝑆 𝑏
EQUATION 21: VERTICAL TAIL SIZING
Given that the tail by definition extends to the upper wing and that the tip chord must
have about the same chord as the root of the upper wing, the range of options for tail AR is
limited. To increase the size of the tail the root chord is modified without changing anything else.
If the tail root chord becomes excessive, then the tail extends above the upper wing. In the final
configuration, the root chord is slightly greater than the tip chord and the tail extends to the
upper wing without going above it.
Once the tail size is determined in this manner, it is checked against the tail size required
when there is one engine non-operational. This set of equations is also give in Airplane Design:
Part II47, but does not end up modifying the final result. In the final configuration, the rudder
deflection is 9.64 degrees for operating with one engine out and the tail size is more than
adequate from a control standpoint. The stability results are analyzed later.
High Lift Devices
The control surface design methodology used for this aircraft is a hybrid of wind tunnel
data and numerical analysis combined with the standard flap chord fraction of 25%. As
mentioned in the airfoil selection section, data was gathered regarding the airfoil change in cl
for single flaps and double slotted fowler flaps; however, this data cannot be directly applied as
it is an airfoil modification rather than a wing modification. As a result, a double slotted fowler

36
flap is used on the aircraft. The aerodynamic analysis software used for this project, Athena
Vortex Lattice (AVL), does not have the code to utilize double slotted fowler flaps and thus a
single pivot flap is applied to the model instead; however, as the software does not include flap
separation behavior, the single flap is extended to 30 degrees to compensate for the reduced
efficiency of this type of flap. Using these criterion, the takeoff and landing conditions were
checked using AVL to ensure that the flaps provided adequate change in CL without requiring
excessive amounts of corrective moment correction.
Control Surfaces
The control surfaces are designed so that failure can be accommodated and does not
prevent the aircraft from making necessary maneuvers. As such, each control surface has two
independent electric actuators and each control method has three distinct control surfaces on
each wing. This yields a high amount of redundancy that makes the possibility of losing
significant control authority extremely unlikely. Redundancy is guaranteed as a minimum of four
and a maximum of eight independent actuators on each wing would have to become
inoperable for loss of control; as there are four wings, as many as 32 independent control
actuators could fail and the aircraft would still be maneuverable. Furthermore, electric
actuators are less prone to failure than their hydraulic counterparts10, resulting in a very safe
aircraft.
As there is no distinct horizontal tail, the elevators and ailerons have been combined into
a single set of flaps, called elevons. This reduces mechanical complexity while only slightly
increasing program complexity. These elevons are positioned such that the smaller surfaces are
further from the neutral point (NP), which ensures that all of the surfaces exert similar pitching
moments. The flower flaps are located in between the elevons. The flap chord fraction is 25% in
order to provide sufficient control authority during takeoff and landing. The final flap layout is
shown in Figure 23.

37
FIGURE 23: CONTROL SURFACE DIAGRAM
Stability and Control
Stability margins between 20% and -10% are acceptable with modern avionics and
control systems. Greater stabilities are possible but inadvisable as takeoff and landing become
prohibitively difficult. Stability derivatives were determined through the use of AVL. This program
slightly disagrees with the spreadsheet analysis as it does not incorporate thrust moments and
therefore the drag amounts are slightly off. For conventional aircraft this is not a concern as the
thrust line is typically very close to the z-cg location; however, the pairing of the box wing and
the requirements of an open-rotor engine preclude a thrust line near the z-cg location.
Fortunately, the stability calculations do not take the moments on the aircraft into account and
thus the stabilities are valid and constant for all flight conditions.

38
The only stability component present within the iterative data mapping loop is the
stability margin. That value is set to be a minimum of -5% for the most aft flight CG. This set
minimum stability is used to determine the desired NP, as seen in Equation 22, and thus the wing
sweep, through an iterative loop.
𝑁𝑃!"# =
𝐶𝐺!"#
𝑐!"#$
+ 𝑆𝑇!"# 𝑐!"#$
EQUATION 22: DESIRED NP GIVEN STMIN
As the sweep has now been set, the maximum stability can be determined though the
more standard equation, as given in Equation 23.
𝑆𝑇 =
𝑁𝑃
𝑐!"#$
−
𝐶𝐺
𝑐!"#$
EQUATION 23: STABILITY MARGIN
In the final configuration, the maximum static margin is determined to be 12%, while the
minimum is set at -5%. Due to the relaxed stability, a stability augmentation system (SAS) is
required. Once this final configuration was determined, the remaining stability derivatives are
taken from AVL and are shown in Table 20 for the cruise CG location for the sizing mission.
TABLE 20: STABILITY DERIVATIVES
CL CY Cl Cm Cn
α 6.583360 -0.000002 -0.000010 -1.362483 -0.000001
β 0.000001 -0.704618 -0.101996 0.000012 0.093945
p' 0.000019 -0.208198 -0.676315 0.000000 -0.034418
q' 8.265108 0.000012 -0.000007 -48.949242 0.000001
r' 0.000000 0.282145 0.084094 -0.000002 -0.062765
As can be seen in Table 20, the aircraft is stable for this CG location, including spirally
stable, which is rare in modern aircraft. For the unstable situations, the SAS described below is
implemented.

39
The SAS controller which is implemented is based on the configuration shown in Figure 24,
where α is the angle of attack and q is the pitch rate. The system is tuned so that the properties
shown in Figure 25 are met.
FIGURE 24: LONGITUDINAL SAS CONTROL39
FIGURE 25: SAS RESPONSE PROPERTIES39
Weight and Balance
The weight and balance of the aircraft is inherent to its flight characteristics. The initial
weight values were extracted from an average of the suggested weight fractions in Airplane

40
Design: Part V50. As the design process progressed, about half of the values were replaced
values given by other equations and most of the remaining values were given weight reduction
fractions due to the use of materials not available when Roskam compiled his data. The final
weight values for the components are given in Table 21. The empty weight of the aircraft is then
compared with the empty weight output by the weight sizing sheet to modify the advanced
material weight correction factor (η), as mentioned in previous sections.
TABLE 21: FINAL COMPONENT WEIGHT VALUES FOR TAKEOFF DURING SIZING
MISSION50
Component Weight
Fraction
Roskam
Weight (lbs)
Reduction
(%)
Known
Weight (lbs)
Final Weight
(lbs)
Front Wing 0.0531 2403.5 0 1837.9 1837.9
Rear Wing 0.0369 1669.1 0 1423.8 1423.8
Tail/Emp 0.03 1357.5 5 -- 1289.6
Engines 0.125 5656.3 15 -- 4807.8
Fuselage 0.1 4525.0 20 -- 3620.0
Nacelle 0.055 2488.8 0 -- 2488.8
Fixed
Equipment
0.145 6561.3 0 6990.0 6990.0
Passengers -- -- 0 15075.0 15075.0
Nose Gear 0.0036 162.9 5 -- 154.8
Main Gear 0.03 1357.5 5 -- 1289.6
Fuel -- -- 0 5399.0 5399.0
Tfo -- -- 0 54.0 54.0
Crew -- -- 0 820.0 820.0
The CG locations in both the x and z directions are then determined using conventional
means. These values are computed for multiple flight conditions representing the extremities of
the flight envelope. CG values are also computed for when the aircraft is fueled and no crew
are present as well as for when the aircraft is completely empty; these values are not important
for CG travel but must be considered when designing the landing gear. The CG excursion
diagram of the final configuration can be seen in Figure 26.

41
FIGURE 26: CG EXCURSION DIAGRAM
The CG location is then used to modify the sweep angle, given the desired minimum
stability margin.
Design Point Selection Iterative Process
The data map is computed through an iterative loop using a VBA Macro Run Microsoft
Excel spreadsheet which incorporates most aspects of the design. An outline of this process is
shown in Figure 27.

42
FIGURE 27: ITERATIVE PROCESS WORKFLOW
As the results are returned, the overall result is iterated until all the variables have
converged to within 0.25 units of their desired value. As the last variable to converge is η, most
of the variables are within 0.005 of their desired value and the L/D values are accurate to within
0.001. Once all of the values have converged, the result is output to the data map and the next
set of parameters are input. Due to the computational time involved in computing data maps,
several maps are used such that they were modified from a large map with low resolution to
smaller maps with high resolution. Once the desired operating range is determined and
additional variables are needed to be set to meet other requirements, the old variables
become set and the new variables are iterated on.
The initial parameters are altitude, cruise speed, climb rate, aspect ratio, and h/b. The
output parameters include fuel burn, takeoff weight, wing sweep, and all of the input
parameters. The results from this data map were shown in Figure 19, Figure 20 and Figure 21 and
indicate that the optimal flight condition is 31,000 feet at Mach 0.68. This may seem counter-
intuitive; however, this cruise condition is enforced by the drag bucket. As the speed decreases
the required cl increases and therefore the operating condition climbs the side of the drag
bucket rather than remaining in the trough, which is where the operating condition lies when the
Mach number is 0.68 at 31,000 ft. Furthermore, the climb speed becomes set to a value of 470
feet per second, as that is the value for least fuel burn that still satisfies the FAA requirements on
climbing at a 3% gradient while simultaneously ensuring that the climb time did not exceed 35
minutes. The climb speed requirement is shown in Figure 28. Lastly, the h/b was set to 0.2 as this is
the maximum value permitted to avoid excessive tail size while providing the lowest fuel burn.

43
FIGURE 28: CLIMB SPEED VS. ALTITUDE
As a note, the ratio between forward and aft wings is important and if denoted as a/f.
The second map included CG travel in order to minimize trim issues. Thus the variables
for this set of maps are: tail forward (the distance from the back of the tail to the back of the tail-
cone), AR, mean chord of the fore wing divided by the aft wing (c (f/a)), the ratio of the sweep
of the forward wing in relation to the aft wing (Λ (f/a)), and differential incidence angles;
altitude, cruise speed, and climb rate are held constant. The difference of chord is used rather
than a difference in area, as the difference in sweep modifies the geometry in such a way that
a chord ratio cannot be easily transferred into the area ratio. As can be seen in Figure 29, the
further forward the tail, the lower the CG travel do to the decreased fuel weight contributions;
increasing the forward chord reduces its effectiveness. In Figure 30, it is seen that increasing the
angle of the forward wing increases CG travel; this is due to the NP movements which
accompany the change in angle of attack. This value cannot be made positive as then the aft
wing would stall before the forward wing. Figure 31 shows that there is a definite trough in which
to find the minimum CG travel for all configurations of this type, while the relative chords effect
the value of the trough.

44
FIGURE 29: TAIL FORWARD ON CG TRAVEL FOR MULTIPLE C (F/A)
FIGURE 30: DIFFERENTIAL ALPHA ON CG FOR MULTIPLE C (F/A)
FIGURE 31: SWEEP (F/A) ON CG TRAVEL FOR MULTIPLE C (F/A)

45
The final set of iterations takes the cruise moment into account. Contributions to the final
moment from all primary components are included. The fuselage and nacelles only contribute
drag while the engines produce thrust and the wings produce both lift and drag. The moments
are determined by multiplying the force by the moment arm. For drag and thrust, the moment
arm is the vertical displacement from the CG-z while for the lift the moment arm is the horizontal
displacement from the neutral point to CG-x. For these calculations, the CG locations used are
for mid-cruise during the sizing mission.
This process begins by determining the drag fractions from the class II drag polar which
yields the fraction of the drag produced by each component. Then these fractions are
multiplied by the total drag at the flight condition of interest to determine the total amount of
drag produced by each component. The moment arms are then calculated using the data
from the CG tables and the drag moments are calculated. The lift and thrust forces are
extracted from the energy constraint system and multiplied by their respective moment arms.
Lastly, the moment contribution inherent to the airfoils is added. These moments are then
summed and saved in the data map. Values for takeoff and landing are also saved to ensure
sufficient control authority. For the final configuration, the moment coefficient is calculated for
the forces which AVL does not consider (i.e. thrust), so that the values for flap and control
surface deflections can be analyzed and produce reasonable results.

46
The final design values output by the iteration process are given in Table 22.
TABLE 22: FINAL DESIGN PARAMETERS
Variable Value
Power (hp) 8183
W/S 81
Time to Climb (min) 34.35
Climb Fuel (lb) 466.7
Cruise Fuel (lb) 3464.4
Descent Fuel (lb) 133.8
Cruise L/D 12.77
Climb L/D 20.57
Descent L/D 13.23
WTO (lb) 45250.28
Vclimb (fps) 470
η 0.91
Average Λ (°) 25.40
Main Gear Location (ft) 57.50
AR 12
Γ (°) 7
h/b 0.2
Cruise Mach 0.68
Tail Forward (ft) 0
Altitude (ft) 31,000
A(f/a) 1.2
Stability Range 17%
Min Stability -5%
Max Stability 12%
CG Travel (ft) 1.190
Cruise Moment (ft lb) 2.095
Taper Ratio 0.2818
Fuselage
The fuselage design process focuses on passenger comfort and minimizing weight and
drag.

47
Sizing
The seating configuration dictates the minimum allowable diameter for the circular cross-
section of the P-65’s fuselage. Circular cross-sections result in a low wetted area, which reduces
friction drag. Eighteen rows of four-abreast seating with two additional seats at the back of the
cabin provides room for 74 passengers. The seats are modeled after a new “slimline” design
that features thinner structure and padding that gives more apparent legroom between the
seats without sacrificing comfort and safety. Each pair of seats is rated to withstand updated ‘g’
loads as specified by FAR Part 2514. Table 23 shows the P-65’s seating arrangement dimensions
compared to that of the CRJ-700. Figure 32 shows the cross-section of the fuselage along with
critical cabin dimensions labeled in inches.
TABLE 23: SEAT CONFIGURATION DIMENSIONS3
Aircraft Seat
Pitch (in)
Seat
Width (in)
Aisle
Width (in)
Aisle Height
(in)
P-65 31 18 19.5 77
CRJ-700 31 17.3 16 72
FIGURE 32. P-65 CABIN CROSS-SECTION
Historical sizing for the data galley, lavatory, wardrobe, and door along with typical
fuselage ratios determine the remaining length of the fuselage. Figure 33 shows the fuselage
length, tailcone length, and tailcone angle.

48
FIGURE 33. OVERALL FUSELAGE DIMENSIONS
The hatched area represents overhead and under floor baggage compartments, and
provides 300 ft3 of cargo volume, which is adequate for a full passenger load37. Figure 34 shows
the final floor plan of the fuselage.
FIGURE 34. FUSELAGE INTEROR CONFIGURATION
Table 24 shows the dimensions selected for the P-65 fuselage, which either match or
improve on the values for comparable jet aircraft in terms of passenger comfort48.
TABLE 24: P-65 FACILITY DIMENSIONS
Facility Dimensions (in.)
Galley (fwd) 46x30
Galley (aft) 48x32
Lavatory (x2) 44x38
Wardrobe 46x30
The fuselage nose incorporates a smooth, cambered ogive shaped body to promote
NLF over a portion of the forward fuselage. Despite the relatively small region experiencing NLF,
the effect helps the P-65 achieve the estimated 25% reduction in skin friction drag19.

49
Noise Reduction
The P-65 fuselage incorporates active and passive noise reducing systems. Lightweight
SOLIMIDE AC-550 foam lines the fuselage between the skin and internal cabin interior panels.
This high-density foam provides thermal and acoustic insulation for the passengers and crew. In
addition, the P-65 features a noise and vibration suppression (NVS) system similar to that found
on the Bombardier Q4006. This system uses microphones installed in the aft portion of the cabin
to detect pulses of air hitting the empennage caused by the rotating engine blades.
Piezoelectric vibration absorbers mounted on the fuselage structure produce pulses to cancel
out the original vibrations, resulting in a quieter and more comfortable experience.
Doors
Passenger access doors, service access doors, and emergency exits are the three types
of doors required for commercial transport aircraft. One 72”x36” passenger access door on the
port side of the aircraft is sufficient for an aircraft carrying 74 people37. Several service doors
enable technicians and ground crews to service important subsystems. Two 31”x31” baggage
doors on the starboard side of the aircraft provide access to the two baggage compartments
forward and aft of the wing box. Exit doors are sized in compliance with FAR Part 25 Sections
807-81314.
Structure
The P-65 implements a semi-monocoque fuselage design where bulkheads, frames, and
stringers provide support to a thin cylindrical outer skin. In order to reduce weight, the fuselage
skin, frames, and stringers are constructed from high modulus, epoxy/resin, carbon fiber
reinforced polymer (CFRP) composites using a variety of fabrication techniques. Carbon
composites have a higher stiffness, significantly higher strength to density ratio, and better
corrosion resistance than comparable aluminum components. Figure 35 shows a color-coded
structural representation of the P-65 with a legend that identifies the highlighted components.

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