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The Pegasus Aerospace Team
Nicholas Robson Georgia Tech BSAE ‘14 Member #: 487875
Kieffer Mil l igan Georgia Tech BSAE ‘14
Coll in Strassburger Georgia Tech BSAE ‘14 Member #: 463147
Airth Burtman Georgia Tech BSAE ‘14
Christ ian Rasmussen Georgia Tech BSAE ‘14
Andrew Norris Georgia Tech BSAE ‘14
Josh Rogers Georgia Tech BSAE ‘14
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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 Alt itude: 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 ..................................................................................................... iiExecutive Summary ...................................................................................................................... iii
Fuel Reductions ......................................................................................................................... iiiNew Technologies .................................................................................................................... iiiJet-Like Experience ................................................................................................................... ivCost and Environmental Friendliness ...................................................................................... iv
Table of Contents .......................................................................................................................... viList of Figures ................................................................................................................................. viiList of Tables ................................................................................................................................... ixNomenclature ................................................................................................................................ xIntroduction ................................................................................................................................... 1Regional Jet Outlook .................................................................................................................... 1
Market Analysis ......................................................................................................................... 2Stakeholder Analysis ................................................................................................................. 5
Existing Designs .............................................................................................................................. 6Configuration Selection ............................................................................................................... 7
Figures of Merit .......................................................................................................................... 7Configuration Selection ......................................................................................................... 15
Technologies ................................................................................................................................ 16Box Wing .................................................................................................................................. 16Propfan .................................................................................................................................... 17Composite Structures ............................................................................................................. 17Natural Laminar Flow ............................................................................................................. 17Digital fly-by-wire .................................................................................................................... 18Solid Oxide Fuel Cell Auxiliary Power Unit ............................................................................ 18Riblets ....................................................................................................................................... 18Spyroid Wingtip ....................................................................................................................... 18Circulation Control Wing ....................................................................................................... 18Technology Compatibility ..................................................................................................... 19
Weight Sizing ................................................................................................................................ 20Mission Profile ........................................................................................................................... 20Weight Estimation ................................................................................................................... 20Payload Weight ...................................................................................................................... 21Mission Fuel Weight ................................................................................................................. 21Gross Takeoff and Empty Weights ........................................................................................ 22Drag Polar (Class I Method) .................................................................................................. 23Drag Polar (Class II Method) .................................................................................................. 25
Constraint Sizing .......................................................................................................................... 26K Determination ...................................................................................................................... 27Constraint Sizing Results ......................................................................................................... 29
Wing Design ................................................................................................................................. 31Planform Selection .................................................................................................................. 31Airfoil Selection ........................................................................................................................ 34Empennage Sizing .................................................................................................................. 34High Lift Devices ...................................................................................................................... 35Control Surfaces ..................................................................................................................... 36Stability and Control ............................................................................................................... 37
Weight and Balance .................................................................................................................. 39Design Point Selection Iterative Process ................................................................................... 41
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Fuselage ....................................................................................................................................... 46Sizing ......................................................................................................................................... 47Noise Reduction ...................................................................................................................... 49Doors ........................................................................................................................................ 49Structure ................................................................................................................................... 49Flight Deck Design .................................................................................................................. 52
Propulsion ..................................................................................................................................... 54Core Selection ........................................................................................................................ 54Number of Blades ................................................................................................................... 54Tip Speed ................................................................................................................................. 55Blade and Disk Power Loading ............................................................................................. 55Blade Sweep ........................................................................................................................... 56Blade Twist and Taper ............................................................................................................ 57Blade Materials ....................................................................................................................... 57Inlet ........................................................................................................................................... 57Results and Summary ............................................................................................................. 58
Landing Gear .............................................................................................................................. 59Configuration .......................................................................................................................... 59Tip-Over and Clearance ........................................................................................................ 59Tires ........................................................................................................................................... 62Retraction Kinematics ............................................................................................................ 62
Structural Consideration ............................................................................................................. 64Materials .................................................................................................................................. 64V-n Diagram ............................................................................................................................ 64
Environmental Impact ................................................................................................................ 65Systems ......................................................................................................................................... 66
ECS ........................................................................................................................................... 66Fuel System .............................................................................................................................. 71Electrical System ..................................................................................................................... 73
Cost Analysis ................................................................................................................................ 74Flyaway Cost ........................................................................................................................... 74O&M Costs ............................................................................................................................... 77Marketing ................................................................................................................................. 79
References ................................................................................................................................... 84
L ist of Figures Figure 1: Air Traffic RPK Growth Through 20245 .......................................................................... 2Figure 2: Air Traffic by Flow5 .......................................................................................................... 3Figure 3: Regional Aircraft Proportion of World Fleets5 ............................................................. 3Figure 4: Average Gasoline Prices, 1991-201258 ........................................................................ 4Figure 5: Global LCC Growth11 .................................................................................................... 4Figure 6: Stakeholder Analysis ...................................................................................................... 5Figure 7: Q40025 and CRJ70024 ..................................................................................................... 6Figure 9: Box Wing Configuration23 ........................................................................................... 10Figure 10: Turboprop21 and Propfan Engines28 ........................................................................ 12Figure 11: Propfan, Turboprop and Turbofan Propulsive Efficiency Plots31 ........................... 13Figure 12: Preliminary Design Configuration ............................................................................ 15Figure 14: Mission Profile ............................................................................................................. 20Figure 15: Weight Regression for Comparable Aircraft29 ....................................................... 23Figure 17: Comparison of Class I and Class II Drag Polar – Cruise ......................................... 26
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Figure 18: Energy Based Constraint Sizing ................................................................................ 31Figure 19: Altitude vs. Fuel Burn for Multiple Mach at AR = 12 ............................................... 32Figure 20: Altitude vs. Fuel Burn for Multiple AR at M = 0.68 ................................................... 33Figure 21: Fuel Weight vs. Sweep Angle for Multiple AR ......................................................... 33Figure 22: Final Wing Configuration .......................................................................................... 33Figure 23: Control Surface Diagram ......................................................................................... 37Figure 24: Longitudinal SAS Control39 ........................................................................................ 39Figure 25: SAS Response Properties39 ........................................................................................ 39Figure 26: CG Excursion Diagram .............................................................................................. 41Figure 27: Iterative Process Workflow ........................................................................................ 42Figure 28: Climb Speed vs. Altitude .......................................................................................... 43Figure 29: Tail Forward on CG Travel for Multiple c (f/a) ........................................................ 44Figure 30: Differential Alpha on CG for Multiple c (f/a) ......................................................... 44Figure 31: Sweep (f/a) on CG Travel for Multiple c (f/a) ........................................................ 44Figure 32. P-65 Cabin Cross-Section ......................................................................................... 47Figure 33. Overall Fuselage Dimensions ................................................................................... 48Figure 34. Fuselage Interor Configuration ................................................................................ 48Figure 35. Fuselage Structure Model ......................................................................................... 50Figure 36. Sheet-Stringer Approximation9 ................................................................................. 51Figure 37. Bay Structure (Dimensions in Inches) ....................................................................... 52Figure 38. P-65 Cockpit Visibility ................................................................................................. 53Figure 39. P-65 Cockpit17 ............................................................................................................ 53Figure 40: Blade Diameter vs. Blade loading ........................................................................... 55Figure 41: Spin Rate vs. Power Loading per blade ................................................................. 56Figure 42: Final Engine Design ................................................................................................... 59Figure 43: Longitudinal Tip-Over Criterion Check47 ................................................................. 60Figure 44: Lateral Tip-Over Criterion47 ....................................................................................... 61Figure 45: Lateral Tip-Over Criterion Check ............................................................................. 61Figure 46: Longitudinal Ground Clearance Check ................................................................. 61Figure 47: Lateral Ground Clearance Check .......................................................................... 62Figure 48: Nose Gear Stick Diagram ......................................................................................... 63Figure 49: Main Gear Stick Diagram ......................................................................................... 63Figure 50: Nose (L) and Main (R) Gear ..................................................................................... 63Figure 51. Material Representation ........................................................................................... 64Figure 52: V-n Diagram ............................................................................................................... 65Figure 53: Environmental impact Compared to Baseline ...................................................... 65Figure 54: Cabin temperature Control Zones .......................................................................... 66Figure 55: Mechanical Piping Schematic (Under Floor) ......................................................... 67Figure 56: Air Distribution Schematic (Above Ceiling) ............................................................ 68Figure 57: Return Air & Exhaust Schematic ............................................................................... 68Figure 58: Air Flow Visualization ................................................................................................. 68Figure 59: Environmental Control System Visualization ........................................................... 69Figure 60: Noise Sensitivity to Flow Velocity .............................................................................. 70Figure 61: Sound Levels at Outlets ............................................................................................ 70Figure 62: Regenerative Noise Reduction Based On Plenum Design ................................... 71Figure 63: Fuel System Schematic ............................................................................................. 73Figure 64: P-65 Wiring Layout ..................................................................................................... 74Figure 65: Pegasus 65 Flyaway Cost Curve .............................................................................. 76Figure 66: Flyaway Cost Breakdown By Individual Element ................................................... 77Figure 67: Flyaway Cost Comparison to Similar Class Aircraft ............................................... 77Figure 68: Breakdown of O&M Costs ........................................................................................ 78Figure 69: Comparison of CASM Costs for Similar Class Aircraft ............................................ 79
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Figure 71: Dimensioned 3 View ................................................................................................. 81
L ist of Tables Figure 1: Air Traffic RPK Growth Through 20245 .......................................................................... 2Figure 2: Air Traffic by Flow5 .......................................................................................................... 3Figure 3: Regional Aircraft Proportion of World Fleets5 ............................................................. 3Figure 4: Average Gasoline Prices, 1991-201258 ........................................................................ 4Figure 5: Global LCC Growth11 .................................................................................................... 4Figure 6: Stakeholder Analysis ...................................................................................................... 5Figure 7: Q40025 and CRJ70024 ..................................................................................................... 6Figure 9: Box Wing Configuration23 ........................................................................................... 10Figure 10: Turboprop21 and Propfan Engines28 ........................................................................ 12Figure 11: Propfan, Turboprop and Turbofan Propulsive Efficiency Plots31 ........................... 13Figure 12: Preliminary Design Configuration ............................................................................ 15Figure 13: Mission Profile ............................................................................................................. 20Figure 14: Weight Regression for Comparable Aircraft29 ....................................................... 23Figure 16: Comparison of Class I and Class II Drag Polar – Cruise ......................................... 26Figure 17: Energy Based Constraint Sizing ................................................................................ 31Figure 18: Altitude vs. Fuel Burn for Multiple Mach at AR = 12 ............................................... 32Figure 19: Altitude vs. Fuel Burn for Multiple AR at M = 0.68 ................................................... 33Figure 20: Fuel Weight vs. Sweep Angle for Multiple AR ......................................................... 33Figure 21: Final Wing Configuration .......................................................................................... 33Figure 22: Control Surface Diagram ......................................................................................... 37Figure 23: Longitudinal SAS Control39 ........................................................................................ 39Figure 24: SAS Response Properties39 ........................................................................................ 39Figure 25: CG Excursion Diagram .............................................................................................. 41Figure 26: Iterative Process Workflow ........................................................................................ 42Figure 27: Climb Speed vs. Altitude .......................................................................................... 43Figure 28: Tail Forward on CG Travel for Multiple c (f/a) ........................................................ 44Figure 29: Differential Alpha on CG for Multiple c (f/a) ......................................................... 44Figure 30: Sweep (f/a) on CG Travel for Multiple c (f/a) ........................................................ 44Figure 31. P-65 Cabin Cross-Section ......................................................................................... 47Figure 32. Overall Fuselage Dimensions ................................................................................... 48Figure 33. Fuselage Interor Configuration ................................................................................ 48Figure 34. Fuselage Structure Model ......................................................................................... 50Figure 35. Sheet-Stringer Approximation9 ................................................................................. 51Figure 36. Bay Structure (Dimensions in Inches) ....................................................................... 52Figure 37. P-65 Cockpit Visibility ................................................................................................. 53Figure 38. P-65 Cockpit17 ............................................................................................................ 53Figure 39: Blade Diameter vs. Blade loading ........................................................................... 55Figure 40: Spin Rate vs. Power Loading per blade ................................................................. 56Figure 41: Final Engine Design ................................................................................................... 59Figure 42: Longitudinal Tip-Over Criterion Check47 ................................................................. 60Figure 43: Lateral Tip-Over Criterion47 ....................................................................................... 61Figure 44: Lateral Tip-Over Criterion Check ............................................................................. 61Figure 45: Longitudinal Ground Clearance Check ................................................................. 61Figure 46: Lateral Ground Clearance Check .......................................................................... 62Figure 47: Nose Gear Stick Diagram ......................................................................................... 63Figure 48: Main Gear Stick Diagram ......................................................................................... 63Figure 49: Nose (L) and Main (R) Gear ..................................................................................... 63
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Figure 50. Material Representation ........................................................................................... 64Figure 51: V-n Diagram ............................................................................................................... 65Figure 52: Environmental impact Compared to Baseline ...................................................... 65Figure 53: Cabin temperature Control Zones .......................................................................... 66Figure 54: Mechanical Piping Schematic (Under Floor) ......................................................... 67Figure 55: Air Distribution Schematic (Above Ceiling) ............................................................ 68Figure 56: Return Air & Exhaust Schematic ............................................................................... 68Figure 57: Air Flow Visualization ................................................................................................. 68Figure 58: Environmental Control System Visualization ........................................................... 69Figure 59: Noise Sensitivity to Flow Velocity .............................................................................. 70Figure 60: Sound Levels at Outlets ............................................................................................ 70Figure 61: Regenerative Noise Reduction Based On Plenum Design ................................... 71Figure 62: Fuel System Schematic ............................................................................................. 73Figure 63: P-65 Wiring Layout ..................................................................................................... 74Figure 64: Pegasus 65 Flyaway Cost Curve .............................................................................. 76Figure 65: Flyaway Cost Breakdown By Individual Element ................................................... 77Figure 66: Flyaway Cost Comparison to Similar Class Aircraft ............................................... 77Figure 67: Breakdown of O&M Costs ........................................................................................ 78Figure 68: Comparison of CASM Costs for Similar Class Aircraft ............................................ 79Figure 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.
3
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.
F IGURE 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.
F IGURE 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.
F IGURE 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.
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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.
Exist ing 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.
F IGURE 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 Alt itude (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 Compatibi l i ty 3 Stal l Characterist ics 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
F IGURE 8: CONVENTIONAL (TOP) 24, BWB
(MID)21 AND FLYING WING DESIGNS (BOT)27
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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.
F IGURE 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 Compatibi l i ty 5 5 3 1 Stal l Characterist ics 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 Compatibi l i ty 5 5 Stal l Characterist ics 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 Fuel/Drag Reduction 3 4 Innovation 3 3 Design Cost 4 3 Airport Compatibi l i ty 5 5 Stal l Characterist ics 3 3 Passenger Perception 2 4 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 Fuel/Drag Reduction 3 3 Innovation 3 3 Design Cost 3 3 Airport Compatibi l i ty 5 5 Stal l Characterist ics 3 3 Passenger Perception 1 4 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 Pul ler Fuel/Drag Reduction 3 3 Innovation 4 3 Design Cost 3 4 Airport Compatibi l i ty 5 5 Stal l Characterist ics 3 3 Passenger Perception 5 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 F ly -by-wire
SOFC APU
TABLE 9: P-65 TECHNOLOGY COMPATIB ILITY MATRIX
O = incompatible X = compatible Box W
ing
Propfan
Com
posites
NLF
Digital Fly-by-w
ire
SOFC
APU
Riblets
Spyroid W
ingtip
CC
W
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 Siz ing 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.
F IGURE 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 (
1𝑀!!
)
EQUATION 1: BREGUET’S ENDURANCE EQUATION
𝑅 =375(𝐿/𝐷)𝜂!
𝑐! ln (
1𝑀!!
)
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.
F IGURE 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!,! =
𝑓𝑆,𝐾 =
1𝜋𝑒𝐴𝑅
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!
=
𝐿12 𝜌!𝑉
!𝑆𝐷
12 𝜌!𝑉
!𝑆
=𝐿𝐷
EQUATION 8: L/D
25
TABLE 14: PHASE SPECIFIC L/D
Phase L/D (-) Cl imb 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 I I Parasit ic 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.
F IGURE 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 I I DRAG POLAR – CRUISE
Constraint Siz ing 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!
𝑉𝑞 𝑆
𝛽 𝑊!"𝐶!,! + 𝐾
𝑁 𝛽! 𝑊!"
𝑞! 𝑆
!
+𝑅!𝑞! 𝑆
+1𝑉𝑑𝑑𝑡
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 − 2𝑑!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
2.83ℎ𝑏
!
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 ∗𝑀0.3
− 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
𝑘1𝑘2𝑘3𝑘4
=1.0370.5711.0372.126
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
Mis
sio
n S
eg
me
nt
Takeoff 0.998 1.000 1.000 ~ 4000 1.1 High Alt itude 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 Ceil ing
0.983 0.361 0.490 94.81 ~ ~
Approach 0.998 1.000 1.000 ~ 4000 1.3 High Alt itude 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
Mis
sio
n S
eg
me
nt
Takeoff ~ 1 0 0 0 0.0184165 High Alt itude Takeoff
~ 1 0 0 0 0.0184165
Climb ~ ~1 0 15.04 0 0.0184166 Cruise ~ 1 0 0 0 0.0187853 Operational Ceil ing
~ ~1 0 15.04 0 0.0184166
Approach ~ 1 0 0 0 0.0184165 High Alt itude 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.
F IGURE 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
Airfoi l c l,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 =𝑁𝑃𝑐!"#$
−𝐶𝐺𝑐!"#$
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.
F IGURE 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 Tai l/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 Tai l Forward (ft) 0 Alt itude (ft) 31,000 A(f/a) 1.2
Stabil ity Range 17%
Min Stabil ity -5%
Max Stabil ity 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)
Ais le Width ( in)
Ais le Height ( in)
P-65 31 18 19.5 77 CRJ-700 31 17.3 16 72
F IGURE 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
Facil ity 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.
50
FIGURE 35. FUSELAGE STRUCTURE MODEL
Frames in the cabin portion of the fuselage are generally spaced every 31 inches in order
to match the seat pitch. This provides each row with an eight-pane window, similar to that
found on the 787, placed in the optimal position to reduce noise and maximize passenger
comfort. Fuselage areas experiencing higher stress concentrations, such as the wing, gear, tail,
and engine attach locations, require additional structure to withstand these loads. The spacing
of the reinforced frames, consisting of back-to-back C-sections, is varied to match the spar,
trunnion, vertical tail, and pylon structure respectively. Sixteen radially spaced continuous and
discontinuous stringers act as stiffeners between frame stations.
Fabrication
Automated fiber and tape placement techniques used in conjunction with pre-
impregnated fiber layup and resin transfer molding increase fabrication efficiency and reduce
tooling costs. A filament-winding process is used to fabricate the nose and empennage due to
their complex curvature. The cabin portion of the fuselage is constructed in two halves that are
bonded and fastened together along the upper and lower seams. Each sub-assembly features
integrally molded stringers, frames, and honeycomb stiffened panels in addition to integrated
sub-structure that is co-cured and co-bonded in an autoclave. The nose, cabin, and
empennage sections are assembled together using sealant and doubler panels to maintain the
integrity of the pressure vessel.
51
Frame and Stringer Sizing
A simplified strength check validates the structural configuration selected from historical
frame and stringer data. The compressive strength of CFRP composites is substantially less than
its strength in tension. As a result, failure due to compressive loads drive the initial structure sizing
process. A 2.5g emergency landing immediately after takeoff generates an approximate
moment of 550,000 ft-lbs in the fuselage at the location of the main gear. The applied load
consists of 2.5 times half of the wing and fuel weight, empennage weight, and engine weight.
The ultimate compressive stress of a quasi-isotropic, high-modulus [0/±45/90](n)s laminate is
approximately 70,000 psi, which is further adjusted by a design reduction factor of 0.5 to
account for the allowable buckling stress between frames57. For circular sections, the average
compressive stress can be taken as 0.67 times the allowable stress, and the approximate resisting
moment arm of the internal couple equals 0.75 times the diameter. Figure 36 shows the sheet-
stringer approximation and the equivalent system assumption.
F IGURE 36. SHEET-STRINGER APPROXIMATION9
Equation 24 shows that the internal resisting moment and allowable compressive stress
(σ) result in a required area moment of inertia (Iyy) for the equivalent stringer section, where the
required radius of the circular stringer approximation relates to the area moment of inertia using
the parallel axis theorem.
52
𝜎!""#$ =𝑀 𝑧2 𝐼!!
,𝑤ℎ𝑒𝑟𝑒 𝐼!! =𝜋4𝑟! + (𝜋𝑟! 𝑧!)
EQUATION 24: COMPRESSIBLE STRESS AND AREA MOMENT OF INERTIA57
The resulting individual stringer area equals 0.63 in2, which is applied to the unsymmetrical
bulb stringer cross-section. Figure 37 shows the typical bay configuration along with the
dimensions of the skin, frames, and stringers.
F IGURE 37. BAY STRUCTURE (DIMENSIONS IN INCHES)
The C-section frames are sized using historical data for comparable aircraft 48. Cutouts in
the web and outer flange allow the continuous stringers to pass through. Bonded shear clips
attach the frames to the stringers at the cutout locations.
Flight Deck Design The P-65’s avionics and cockpit design focuses on creating a familiar environment for
current pilots that effectively communicates critical information about the aircraft. Figure 38
shows that the cockpit visibility complies with the +45/-15 degree requirement specified by the
FAR Part 2514.
53
FIGURE 38. P-65 COCKPIT VISIBILITY
The layout resembles that of the CRJ-700 in order to facilitate airline pilots’ transition from
current regional jet aircraft to the P-65. Figure 39 shows the modified CRJ-700 cockpit design
featuring the Garmin G5000 flight deck system.
FIGURE 39. P-65 COCKPIT17
54
Propulsion
Core Selection In order to select a core for the engine, a few different options are considered. These
cores are near the required power that is required based on the constraint sizing, and are
tweaked and improved upon as necessary in order to meet the mission requirements of the P-65.
The cores considered can be found in Table 25.
TABLE 25: ENGINE CORE DATA
Engine Model Weight ( lb) Power (shp) Power/Weight Pratt and Whitney PW127G40 929 3058 3.29 RR(All ison) T5645 1940 4350 2.24 RR(All ison) AE210044 1727 4637 2.69 Pratt and Whitney PW15040 929 5071 5.46
The PW 127G and the PW150A are outside of the required power for each engine
(approximately 4,200 horsepower). However, they are both derivatives of the PW100, which has
a broad spectrum of available engines. An engine could easily be modified to fit the mission
profile. The T56 and the AE2100 are the two engines that are used on the Lockheed C130
Hercules, a military transport. The PW150A is used on the Bombardier Q400, a baseline aircraft
for this design.
The PW100 series is selected due to its heavy use in the regional aircraft industry and its
superior power-to-weight ratio. The increase in power-to-weight ratio allows the aircraft to be
lighter, and thus more efficient. Also, considering the light weight of the aircraft, it helps with the
weight and balance.
Number of Blades Van Zante and Wernet54 use a configuration comprising of 12 forward blades and 10 aft
blades. This configuration is used for this design as well. This number of blades provides good
numbers for the power loading per blade, which is seen in the following section.
55
Tip Speed Reynolds, Riffel, and Ludemann42 use 750 feet per second as the tip speed for their
design. This tip speed is not too high as to cause many problems with drag divergence or noise.
Using a higher tip speed would cause more of these issues at cruise.
Blade and Disk Power Loading The shaft power from the core is split into two separate, counter-rotating disks, each
having an equal amount of power. Since each core produces approximately 4,200 shaft
horsepower (shp), the approximate shaft power going to each disk is approximately 2,100 shp.
Using typical blade loadings for turboprops from Airplane Design: Part II47, a range of
acceptable blade loadings was determined, and a trade study was performed in order to see
how the blade loading affected the disk rotational speed. Equation 25 is used to calculate
blade loading. The results of this trade study can be seen in Figure 40 and Figure 41.
𝑃!" =4𝑃!"#𝜋𝑛!𝐷!!
EQUATION 25: BLADE LOADING47
FIGURE 40: BLADE DIAMETER VS. BLADE LOADING
56
FIGURE 41: SPIN RATE VS. POWER LOADING PER BLADE
It can be seen from the plots that both relationships are nearly linear within the range of
power loadings and are within the reasonable range presented by Roskam for regional
turboprops. Since the spin rates are all relatively low within this range, the highest loading within
the range is selected so that the propeller diameter is minimized, allowing for easier integration
and reduced engine weight. The prop spin rate is still lower than what is typically expected for
turboprops, but this is acceptable due to the fact that the P-65 cruises at a higher Mach
number. In order to keep the tip speed low at a higher cruising Mach, the propeller must spin
more slowly. Transonic effects on the tips might increase cabin noise and decrease propulsive
efficiency. The final diameter of the propeller is 6.36 feet while the final spin rate is roughly 754
RPM.
Blade Sweep Sweep is utilized on the blades in order to avoid transonic effects on the blades. The
blades are swept back. In order to calculate the sweep required, the Mach number normal to
the blade is chosen. This Mach number is set to be low enough to avoid transonic effects, and is
not to be exceeded at any point on the blade. The Mach number should not be greater than
0.72 since the blades are fairly thin. This Mach number is first experienced 1.04 feet from the
57
blade root, which is where the sweep begins. It increases linearly until the tip, where the sweep is
17.57 degrees, given the tip speed of 750 ft/sec or Mach 0.76 at cruise; the tip sweep is finalized
at 17.6 degrees for the forward blades.
For the aft blades, the tip sweep is increased to account for the acceleration imparted
on the flow by the first stage. Using the power from the first stage, it is estimated that the velocity
going into the second stage is 187 feet per second faster than the freestream velocity
experienced by the first stage. As a result, the blade tip sweep for the aft blade is 40.25 degrees.
Blade Twist and Taper In order to avoid stalling the blades near the tips, the blades have linear twist throughout
the span. The blades are variable pitch, but the relative wind at the tip is vastly different from the
relative wind at the root. The difference between the relative wind vectors is approximately 21
degrees for the forward blades and 20.75 degrees for the aft blades. These angles are
calculated based on the velocity contributions from the freestream direction and the direction
in which the blades are moving when spinning. Taper is also considered. Due to tip losses, it is
undesirable to have a large blade chord at the tip of the wing. The taper ratio of the wing is 0.2
in order to avoid these undesirable effects.
Blade Materials The blade materials used by Hamilton Sustrand are aluminum spars with a carbon
composite skin and a foam or honeycomb core42 so the honeycomb core is selected for this
engine due to its high strength to weight ratio. Short blades also require less structural stiffness
than traditional turboprop blades.
Inlet The inlet design is guided by Hill and Peterson’s Mechanics and Thermodynamics of
Propulsion18. The authors provide descriptions on how to use the inlet to efficiently slow down the
flow before it enters the compressor. Conservation of mass is the principle by which the flow is
slowed. By increasing the area through which the flow passes, the flow is slowed down by 45
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percent. There are trade-offs associated with this deceleration. As the flow slows, it creates an
adverse pressure gradient through which the flow resists passing. Excessive adverse pressure
gradients can lead to flow separation, stagnation pressure losses, and spillage from the engine.
In order to avoid these effects, Hill and Peterson suggest that the entrance to the compressor be
no more than 30 percent larger than the inlet. In this design, the compressor entrance is 27
percent. As was stated previously, the velocity of the flow decreases by 45 percent through the
inlet; the Mach number entering the inlet is Mach 0.374.
Results and Summary The final engine design is summarized in Table 26 and shown in Figure 42.
TABLE 26: SUMMARY OF ENGINE PARAMETERS
Parameter Value Disk Area (ft2) 33.2 Prop Diameter (ft) 6.5 Blade Power Loading (hp/ft2) 6.0 Propeller Spin Rate (RPM) 754 Forward Blade Max Sweep (°) 17.57 Aft Blade Max Sweep (°) 40.25 Max Power (hp) 4200
59
FIGURE 42: FINAL ENGINE DESIGN
Landing Gear
Configuration Due to the speed and landing requirements of the P-65, it has retractable, tricycle
landing gear.
Tip-Over and Clearance To determine whether the ground clearance and tip-over criterion are satisfied, an initial
landing gear strut disposition was selected based on the weight and balance data for the
design. The disposition information is listed in Table 27.
TABLE 27: LANDING GEAR DISPOSITION
Disposit ion Specif ication Nose Gear Location 10 ft. from nose Main Gear Location 57 ft. from nose Main Gear Width 6 ft. from fuselage centerline Ground Clearance 3.5 ft.
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Longitudinal Tip-Over To prevent longitudinal tip-over for tricycle landing gear, the main gear is located behind
the most aft CG such that a line between the main gear and the most aft CG creates at least a
15° angle with the vertical axis47. Figure 43 shows that for the chosen main gear displacement,
this criterion is satisfied.
F IGURE 43: LONGITUDINAL TIP-OVER CRITERION CHECK47
Lateral Tip-Over To prevent lateral tip-over, the criterion illustrated in Figure 44 should be satisfied. This
criterion is dependent on the angle ψ and necessitates that this value should be no greater than
55°47. Figure 45 shows that for the landing gear strut dispositions chosen; this angle is equal to
52.9° meaning that this criterion is satisfied.
61
FIGURE 44: LATERAL TIP-OVER CRITERION47
F IGURE 45: LATERAL TIP-OVER CRITERION CHECK
Ground Clearance The longitudinal ground clearance criterion for tricycle gear requires that the angle θ
should be at least 15° to ensure that the tail end of the aircraft will not hit the ground during
takeoff47. Figure 46 shows that for the landing gear placement chosen the criterion is satisfied.
FIGURE 46: LONGITUDINAL GROUND CLEARANCE CHECK
62
To ensure that the wings or other components attached to the wings do not come in
contact with the ground, the angle φ as defined in the drawing must be greater than 5°47. Figure
47 shows that for the landing gear placement chosen, this criterion is satisfied.
FIGURE 47: LATERAL GROUND CLEARANCE CHECK
Tires Using the method prescribed in Airplane Design: Parts II47 and IV49, Table 28 gives the
necessary tire information for the nose and main gear.
TABLE 28: SELECTED TIRE INFORMATION
Parameter Nose Main Diameter ( in) 23 34 Width ( in) 7 11 Max Loading (lbs) 6500 16100 Max Speed (mph) 210 200 Unloaded Inflation Pressure (psi) 110 145
Retraction Kinematics The retraction stick diagram for the nose gear is located in Figure 48 and the stick
diagram for the main gear is located in Figure 49. Figure 50 shows the resultant gear.
63
FIGURE 48: NOSE GEAR STICK DIAGRAM
FIGURE 49: MAIN GEAR STICK DIAGRAM
FIGURE 50: NOSE (L) AND MAIN (R) GEAR
64
Structural Consideration
Materials
Table 29 shows the properties for the three main structural materials used in the P-65.
Figure 51 gives a visual representation of the structural materials.
TABLE 29. P-65 MATERIAL PROPERTIES33
Parameter Carbon/Epoxy Fiber
Composite
Al-7075 TI-6Al-4V
Density ( lb/in3) 0.0614 0.102 0.16
Tensi le Strength/Yield (ksi) 129 15 138 Tensi le Strength/Density (103 in) 2101 147 863 Compressive Strength/Yield (ksi) 74 15 141 Modulus (ksi) 32000 10400 16500
FIGURE 51. MATERIAL REPRESENTATION
V-n Diagram A V-n diagram is put together with the guidance of FAR Part 2514 regulations and the
results are seen in Figure 52. All airspeed and load requirements are met per the RFP.
65
FIGURE 52: V-N DIAGRAM
Environmental Impact Due to the P-65’s reduced fuel consumption, the environmental impact of the aircraft is
significantly reduced. The yearly carbon dioxide emission of the P-65 is compared to baselines
for both regional turboprop and turbofan aircraft in Figure 53.
FIGURE 53: ENVIRONMENTAL IMPACT COMPARED TO BASELINE
66
Systems Due to the focus of the “jet-like experience” and the unique configuration, the
environmental control system (ECS) and the fuel system are focused on in this section. The
electrical system is also addressed
ECS Pegasus Aerospace’s consumer-centric design mentality made the “jet-like experience”
of the P65 a priority. Among the many design objectives for the cabin are a reduced apparent
cabin altitude, reduced noise pollution from the ECS and propfan engines, and cleaner,
healthier air for the P65’s passengers. This necessitates a complete ECS design.
Design FAR Part 25.831 necessitates a fresh air mass flow rate of 0.55 lbs/min per passenger.
The ECS in the P65 is a no-bleed-air system. While most aircraft use conditioned air that is
bled off from the compressor of the aircraft’s engine, the P-65 uses only ram air, and HEPA
filtered recirculated air. The no-bleed system has substantial benefits; because no air is bled off
the engine, the thermal cycle of the engine becomes more efficient, and engine SFC is reduced
on the order of 3-4%. Additionally, the elimination of bleed air simplifies the ECS, simplifying
ground maintenance, and also reduces material needs and hence aircraft weight.
The P65 cabin has three temperature control zones, shown in Figure 54, that consolidate
the areas within the cabin that have similar heat load characteristics. Zone 1 encompasses the
cockpit, zone 2 contains aft and forward galleys, and zone 3 contains the passenger seating
area.
FIGURE 54: CABIN TEMPERATURE CONTROL ZONES
67
Schematics for the ECS are shown in Figure 55, Figure 56 and Figure 57. The no-bleed ECS
system uses ducted free stream outside air, and splits into a cooling line and heating line after a
flow control valve reduces the inlet flow velocity and pressure. Untreated outside air is used to
cool the APU and cockpit avionics, as shown in the mechanical piping and equipment
schematic in Figure 55. Air that serves the cabin is filtered to remove ozone gas and can either
go through a cooling loop for ground based cabin conditioning or skip the cooling loop for
normal operation. The cooling loop has standard components of heat exchanger (HX),
expansion valve, compressor, and water remover.
F IGURE 55: MECHANICAL PIPING SCHEMATIC (UNDER FLOOR)
The filtered outside air is then piped upward to the above ceiling air distribution shown in
Figure 56, where volumetric flow is controlled using a series of venture valves. The air is then
conditioned to the appropriate temperature through 10-stage heating based upon the
temperature control zone commanded temperatures.
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FIGURE 56: AIR DISTRIBUTION SCHEMATIC (ABOVE CEILING)
FIGURE 57: RETURN AIR & EXHAUST SCHEMATIC
A cabin flow visualization is shown in Figure 58. Return air plenums are consolidated and
located at the flanks of the aisle as shown in Figure 57. The location of the return air plenums are
functional, as the tendency of the cabin airflow is to move from the edges of the cabin toward
the aisles, creating a subtle “sweeping” effect. This will aid ground cleaning crew, and reduce
aircraft turnaround time. A visualization of the ECS is shown in Figure 59.
FIGURE 58: AIR FLOW VISUALIZATION
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FIGURE 59: ENVIRONMENTAL CONTROL SYSTEM VISUALIZATION
ECS Characterist ics & Cabin Noise Reduction Cabin noise is minimized using two main strategies: relocation of equipment, and
minimization of regenerative noise. All mechanical equipment (terminal units, dampers, etc.) is
positioned away from the primary passenger cabin area, as shown in Figure 56, with the main
cabin Venturi valves positioned aft of the rearmost cabin wall. Minimization of regenerative
noise within the air distribution is achieved using a plenum for the main and individual use
passenger lines; this increases duct cross sectional area, and decreases flow velocity. The
plenum ducting is displayed in Figure 56 and Figure 58. Duct sizing, volumetric flow and
velocities are shown in Table 30.
TABLE 30: DUCT SIZING, FLOW VELOCITIES
Line Duct Dim.
ACross-sec. AOutlet CFMReq VOutlet VDuct
Zone 1 ADIST Line 5.5 x 5.5
0.2
0.8
80.0
95.2
190.4
Zone 2 ADIST Lines
4.5 x 3
0.2
0.4
120.0
285.6
285.6
Zone 3 Plenum 40 x 5.5 1.5 4.8 1258.0 259.8 823.4 Zone 3 Individual Use
24 x 3 0.5 0.4 222.0 550.0 222.0
Regenerative noise is a function of flow velocity because increased velocities lead to
increased in-duct turbulence. The positive sensitivity of noise to volumetric flow velocity is shown
70
in Figure 60. The Noise Criteria (NC) is a measure of acceptable noise conditions, ranging from 10
(very quiet) to 55 (very noisy)4.
F IGURE 60: NOISE SENSITIVITY TO FLOW VELOCITY
Within the passenger cabin, the contribution to acoustic levels from the air distribution
outlets is shown in Figure 61 (note that frequency ranges of interest are 1000Hz – 8000Hz4). The
personal distribution units have larger sound power levels than does the main distribution as a
function of the reduced nozzle size at the distribution line’s termination. Total acoustic levels
were optimized by balancing the required volumetric flow between the main plenum (85%) and
individual lines (15%), as seen in Table 30.
FIGURE 61: SOUND LEVELS AT OUTLETS
The primary reduction in cabin noise is from the use of a plenum for the main distribution.
The benefits of the plenum’s large cross sectional area are shown in Figure 62, and are
71
compared to a traditional ECS ducting system4. Note that the plenum, when compared to a
traditional ECS, is nearly inaudible. This is a key feature in the P-65, a level of cabin comfort that
not only rivals a jet-like experience, but exceeds it.
F IGURE 62: REGENERATIVE NOISE REDUCTION BASED ON PLENUM DESIGN
Fuel System As the P-65 has a unique configuration, the fuel system is a subsystem of particular
interest. Fuel is stored in both wings, as neither wing is large enough to accommodate all the
fuel. Equation 26 gives the equation for fuel volume available, which is a function of root
thickness ratio ((t/c)r) and the ratio of thickness ratio at the tip to the thickness ratio at the tip
(τw). For the P-65, the total fuel volume is 167.9 ft3 or 8,592.6 pounds of Jet-A fuel. This shows that
the total wing area has more than enough room to accommodate the fuel necessary to power
the P-65. This also allows for some flexibility in the placement of fuel.
𝑉!" = 0.54(𝑆!/𝑏) 𝑡/𝑐 !{(1 + 𝜆!𝜏!!/! + 𝜆!!𝜏!)/(1 + 𝜆!)!}
EQUATION 26: FUEL VOLUME AVAILABLE47
While though having multiple fuel tanks can introduce some complication, especially
with the vertical distance between the front wing and the engines, these two separate fuel
tanks offer an advantage in the ability to control inflight CG, by having a fuel burn schedule.
Additionally, the fuel tank is designed to reduce overall aircraft CG excursion. This is done by
placing 57.3% of the total fuel in the front wing, even though the front wing represents 59% of the
72
total wing area. This allows for zero CG travel during inflight fuel burn, if the fuel schedule is such
that the fuel is consumed from each fuel tank at a rate proportional to its capacity. Table 31
gives a summary of the front and aft fuel capacities for the maximum fuel loading condition,
assuming that Jet-A fuel with a density of 6.84 pounds per gallon16.
TABLE 31: FRONT AND AFT WING FUEL CAPACITIES
Fuel Tank Fuel Weight ( lbs) Fuel Volume (gal) Front Wing 3096.3 452.9 Aft Wing 2302.7 336.6
Figure 63 shows a basic fuel system schematic. Each tank is separated into four individual
compartments, with baffle valves a part of the walls between each compartment. One-way
baffle valves in both wing tanks allow fuel to move inboard but not outboard so as to prevent
large CG travel during large pitch excursions. The front wing has a dihedral so gravity naturally
draws the fuel inboard to the bottom of the front wing tank, through the baffle valves. The aft
wing has anhedral so scavenge pumps are used to draw fuel inboard. Both front and aft wing
tanks are equipped with collector cells that are kept full at all times by scavenge pumps. High
powered fuel pumps are used to transfer fuel from the wing tanks to the engine. The fuel system
is designed so that fuel can be transferred both between left and right tanks and between front
and aft tanks, if necessary. The fuel line from the front tank runs horizontally from the wing until it
rises to meet the fuel line from the aft tank and then the two join to form one line that runs to the
respective engine. There are two fuel lines running through the fuselage, one for each engine.
The final high-level component of the fuel tank is the vent box, which ensures that air does not
remain in the fuel tanks and cause issues with over-pressurization.
73
FIGURE 63: FUEL SYSTEM SCHEMATIC
Electrical System A general wiring layout for the major electrical lines is designed for the fuselage of the P-
65 in a 2D, top-down view, and can be seen in Figure 64. There are two external power ports
that allow access from either the nose, or tail of the aircraft in order to aid with power needs
while grounded. As a safety precaution, the primary line carrying power from the engines to the
rest of the aircraft is duplicated, and the two lines kept separate on opposite sides of the
fuselage. This redundancy helps to prevent catastrophic loss of power (and control) to the
cockpit if either of the lines are damaged. These two lines have been positioned 36 inches off
center, and are isolated outside of the main structural supports for the fuselage while still being
close enough to the edge that maintenance is not restricted. The lines branch out to the
forward wings from the lower portion of the fuselage, and provide power to the electric
actuators in the wing. Located in the upper wings, the two lines join together at the top of the
vertical tail, providing power to the electric actuators in the wing and tail, and the line travels
down the tail to where it rejoins with the lines coming from the engine pylons. To further add to
the safety of the aircraft, two APUs are used. The first is positioned near the cockpit and the
second is located near the tail and engines. The APUs provide power to the control systems
during the startup procedure, and can also help in supplying power during an inflight power-loss
74
situation. In order to reduce the weight that the reinforced power cables add, remote power
distribution units are positioned along the primary power lines that branch out to nearby
electrical devices with light-weight wire.
F IGURE 64: P-65 WIRING LAYOUT
Cost Analysis There are two different items to look at when performing a cost analysis: the flyaway
cost, or the cost to produce a single aircraft and deliver it to the customer, and the operability
and maintenance (O&M) cost, which is the lifetime cost set over a certain service period, 15
years in this case. The RFP states that the cost of the aircraft should be less than that of a similar
class turboprop and significantly less than that of a similar class regional jet. Therefore this design
seeks to offer cost advantages for both the flyaway cost and the O&M cost. As a note, all cost
numbers are given in 2013 USD.
Flyaway Cost Flyaway cost is determined using the Cost Estimating Relationships (CER) given by
Nicolai35. Flyaway cost is a function of empty weight, maximum cruise velocity and number of
aircraft produced. Additionally, as the structure of the aircraft consists of approximately 50%
75
carbon composite material, difficulty factors are estimated as the CERs are derived for
aluminum aircraft structures only. Difficulty factors are determined using guidance from the
Research and Development (RAND) Corporation42. The RFP requires 400 production aircraft for
the analysis and it is assumed that 5 test aircraft are used in certification. There are eight
discreet cost elements used for this high-level analysis:
• Airframe Engineering: This includes the costs to do the engineering work necessary to
generate a new design, such as research and development as well as testing. The
engineering rate is $120 per hour35 and is assigned a difficulty factor of 1.242. This yields a
total engineering cost of $2.49M per production aircraft.
• Developmental Support: This refers to the nonrecurring cost that is used to support
engineering effort in the design phase. This includes hardware and parts used for testing
and research. Since it is directly related to engineering, it is also assigned a difficulty
factor of 1.242. It has a cost of $0.37M per production aircraft.
• Flight Test: This covers the costs of an effective flight test program that is used to validate
the design of the aircraft. It is assigned a difficulty factor of 1.242 and has a cost of $0.19M
per production aircraft.
• Tooling: This includes the costs relating to the tools and riggings that are used in the
manufacturing of the aircraft. New materials have a greater effect on tooling costs, so it
is given a difficulty factor of 1.442. Tooling costs $127 per hour35, for a cost of $2.11M per
production aircraft.
• Manufacturing Labor: This includes the costs relating to the labor necessary for the
fabrication and assembly of the aircraft. The labor rate is $112 per hour35 and there is
assumed to be no additional difficulty factor for a cost of $6.23M per production aircraft.
• Quality Control: This includes the costs related to assuring the quality of the
manufacturing process and the final product. The quality control rate is $101 per hour35
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and given a difficulty factor of 1.542, due to its importance in the manufacturing process.
This produces a cost of $1.33M per production aircraft.
• Manufacturing Material: This is an important cost element as this design utilizes carbon
composites, along with a 1.8 difficulty factor42. The total material costs are $5.37M per
production aircraft.
• Engines and Avionics: The engines and avionics prices are estimated using existing
design prices as a guideline. The two engines per aircraft cost $2.76M total and the
avionics system is estimated to cost $3.53M per aircraft.
From these subsets the total aircraft flyaway cost is $24.4M. Figure 65 shows the flyaway
cost as a function of units produced, as well as the P-65 cost. Figure 66 shows the flyaway cost
breakdown.
FIGURE 65: PEGASUS 65 FLYAWAY COST CURVE
77
FIGURE 66: FLYAWAY COST BREAKDOWN BY INDIVIDUAL ELEMENT
Using the same analysis, the P-65 production cost is compared to the unit cost for both
the Q400 and the CRJ700. Figure 67 gives a visual comparison of flyaway cost for the three
aircraft.
F IGURE 67: FLYAWAY COST COMPARISON TO SIMILAR CLASS AIRCRAFT
O&M Costs O&M costs are estimated using the guidance of Raymer40. An important cost parameter
for regional jet parameters is that of Cost per Available Seat Mile (CASM). This is determined by
calculating operating costs for a given time period, then dividing the cost by the product of the
miles traveled and the seats available. The typical daily mission of the P-65 consists of five 400 nm
missions and one 1600 nm mission, which takes up 11.1 hours and 4,140 miles. The average
aircraft is assumed to be active 300 days of the year. O&M cost is a factor of the following
elements:
• Fuel: This aircraft uses Jet-A fuel, which has a density of 6.84 lbs/gal16. Fuel is assumed to
cost $5.50 per gallon11. The aircraft burns 626.8 gallons of fuel per day, for a yearly cost of
$1.97M or a CASM of $0.021.
• Maintenance: It is estimated that for every flight hour, 10 hours are needed to maintain
the aircraft. The maintenance rate is assumed to be $40 per hour for a yearly cost of
$1.33M for a CASM of $0.014.
78
• Crew: The crew as a total is estimated to earn $669 per hour. It is estimated that in
addition to the 11.1 hours that are spent in flight, the crew accumulates an additional 3
hours a day in non-flying block time. This yields a yearly cost of $2.83M and a CASM of
$0.030.
• Material Costs: Using the equation provided by Raymer, material costs necessary for the
maintenance of the aircraft total $355 per flight hour, for a yearly cost of $1.18M and a
CASM of $0.012.
In total, O&M costs are $7.31M per year, or $109M over the 15 year lift period. The total
CASM is $0.078.
F IGURE 68: BREAKDOWN OF O&M COSTS
The same O&M analysis is applied to the Q400 and the CRJ700 for a means of
comparison. Table 32 and Figure 69 show that the O&M costs for the Pegasus 65 offer a
significant advantage over similar class aircraft.
TABLE 32: O&M COST COMPARISONS BETWEEN P-65 AND SIMILAR CLASS AIRCRAFT
Cost Pegasus 65 Q400 CRJ700 15 Year O&M Cost (Mil l ions 2013 USD)
109.6
153.1
129.2
79
FIGURE 69: COMPARISON OF CASM COSTS FOR SIMILAR CLASS AIRCRAFT
Marketing Any attempt at marketing must focus on the strengths of the aircraft. These advantages
can be communicated on multiple levels. For the airline companies that are looking to purchase
additional regional aircraft for their fleet, the initial flyaway costs and yearly O&M costs are
appealing when compared to existing aircraft. They can then pass these reduced costs on to
their customers, who are looking for relief from rising ticket prices. Therefore a marketing
campaign would seek to target the airline customers themselves by showing them statistics
about the rising fuel prices that highly contribute to increased ticket prices. Additionally, the
marketing campaign would emphasize its environmental impact and attempt to gain an
endorsement from the Environmental Protection Agency (EPA), which would go on a long way
in a more eco-conscious environment. Government incentives given to companies who use
environmentally friendly engineering solutions would be highlighted, from the aircraft industry to
the energy industry. These benefits show the trend of the current government towards
incentivizing green tech.
The marketing campaign would also seek to mitigate and indeed take head on the
potential aversions that could arise from concerns over a brand new design as well as existing
stigmas relating to turboprops. The public will be kept informed of the certification process. The
innovation of the box wing will be highlighted at several airshows, in hopes of getting the public
80
excited about the design. The safety of the design and the engines will be clearly established.
Focus groups will experience the noise reduction technology of the P-65 firsthand to combat the
association of turboprops with noisiness. In order to combat the perception that turboprops are
an older, less advanced technology, the P-65 will be marketed as “the green future of air
travel.” In this marketing campaign, each of the technologies will be highlighted briefly, along
with their benefits to efficiency and comfort. A trendy advertising campaign, shown in Figure 70,
will help with the public perception of the P-65.
1
FIGURE 70: ADVERTISEMENT FOR P-65 MARKETING CAMPAIGN
81
Conclusion
A dimensioned 3-View is shown in Figure 71, and a compliance compatibility matrix can
be seen in Table 33, showing that Pegasus Aerospace meets and exceeds the requirements of
the RFP in weight and fuel reduction. Additionally, a fuel reduction waterfall diagram is shown in
Figure 72 to show requirement traceability.
The Pegasus Aerospace P-65 is a revolutionary regional turboprop that combines
outstanding fuel efficiency with exceptional passenger comfort, while providing a much needed
innovative design and aesthetic to combat the stigma of use of turboprop aircraft. Pegasus
Aerospace is proud to present the P-65 aircraft to the AIAA – an aircraft engineered from top to
bottom with practicality and passenger experience in mind.
F IGURE 71: DIMENSIONED 3 VIEW
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FIGURE 72: FUEL REDUCTION WATERFALL DIAGRAM
TABLE 33: AIAA RFP COMPLIANCE MATRIX
Category Requirement Compliance P-65
Aircraft Capacity The aircraft shall have a capacity of 75 +/- 2 passengers.
Full 18 rows of four-abreast seating with two additional seats at the back of the cabin provides room for 74 passengers
Economic Miss ion Range and Payload
Economic mission range shall be 400 nmi and the payload shall be the full passenger load, assuming a passenger with baggage weighs 225 lb.
Full The P-65 is capable of completing the economic mission range of 400 nmi while carrying a full passenger load of 74 passengers
Design Miss ion Range and Payload
Design mission range shall be 1,600 nmi with a payload of 67 passengers and with appropriate fuel reserves
Full The P-65 was designed to complete a mission with a range of 1,600 nmi at a 90% load factor while also carrying the proper amount of reserve fuel.
Takeoff and Landing The aircraft shall Full The P-65 is designed to be able to
31.23%
35.4%
30.0%2.30% 0.27% 0.80%
100%
0
20
40
60
80
100
120
P‐65 NLF&WingConfigura=on
Propulsive CompositeMaterials
SOFUAPU Riblets CRJ700
Percen
t Fue
l Use vs. Com
pe3tor (%
)
Technologiies Used
Fuel Reduc3on Waterfall Diagram (vs. CRJ700)
83
be able to use 4,500 ft or shorter runways at maximum takeoff weight (MTOW), assuming sea-level conditions
operate from a 4,000 ft runway at MTOW and standard sea-level conditions.
Non-Standard Takeoff and Landing
The aircraft shall be able to use 8,000 ft or shorter runways at 7,800 ft altitude and 85°F at 80% of the MTOW
Full The P-65 is designed to be able to operate from an 8,000 ft runway at 7,800 ft altitude and 85°F at 80% of the MTOW.
Cruise Speed Cruise Mach number shall be no lower than 0.62 and no higher than 0.68
Full The designed cruise Mach number for the P-65 is 0.68.
Cruise Alt itude The maximum cruise altitude shall be no lower than 25,000 ft and no higher than 31,000 ft
Full The optimal flight condition for the design cruise Mach number of the P-65 occurs at 31,000 ft
Fuel Consumption The aircraft shall demonstrate a fuel consumption that is at least a 65% reduction from a currently operating regional jet of similar seat capacity on a per-seat-nautical mile basis. It should also consume less fuel than currently available turboprop aircraft of similar capacity.
Full The P-65 achieve a fuel consumption reduction of 68.8% off of the baseline.
Cost The predicted cost acquisition and operating costs for the aircraft over a 15 year period for a production run of 400 aircraft shall
Full The flyaway cost of the P-65 is $7.4 million less than the CRJ700 and $0.5 million less than the Q400. The operations and maintenance costs of the P-65 over a 15 year period is $19.6 million less than the CRJ700 and $43.5 million less than the Q400
84
be substantially less than current regional jets and turboprop aircraft of similar capacity
Passenger Comfort/Acceptance
The aircraft shall provide a “jet-like” experience for passengers
Full The P-65 is capable of cruising at a ‘jet-like’ speed of Mach 0.68 to give passengers the speed they desire
Noise Levels Interior cabin noise levels shall be less than that of currently available turboprop aircraft of similar capacity
Full Cabin noise levels in the P-65 are 35.3% lower than comparable aircraft due its to active and passive noise cancellation technologies and plenum air distribution
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