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NASA Ultra Efficient Commercial Airliner ChallengeUniversity of California - Davis
Faculty Advisor: C. P. van DamUndergraduate Level
Team: Joukowski Analysis GroupBruno Matsui, Kilian Ginnell, Adriana Henriquez, Robert Woo, Mengsu Yang
June 15, 2018
Abstract
The weight, sizing, configuration, propulsion, aerodynamics, estimated performance, and projected costare proposed for the XP-1, a middle-of-market small widebody passenger transport aircraft for the 2045 timeframe which achieves the same mission profile as the Boeing 757-300.
The XP-1 passenger transport aircraft leverages advancements in computational density, composite struc-tures, transonic aerodynamics, and hybrid-electric propulsion technology along with a variety of adapted systemsincluding active laminar flow, ultra-sonic de-icing, and cryogenic cooling. The aircraft also features advancedcontrol systems which constantly adapt to changing flight conditions and an elliptical fuselage cross sectionto maximize usable passenger volume per unit of wetted area. Ultimately, by reducing weight and drag andimproving TSFC, the XP-1 attains a 62% decrease in fuel burn per passenger per mile over the baseline Boeing757-300 while achieving a range of 3,500 nm with a full payload of 279 passengers and a maximum takeoffweight that is 37% lower than the baseline. The XP-1 is designed to meet FAR Part 25 airworthiness standardsand be compatible with existing airport infrastructure.
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NomenclatureALF Active Laminar Flow
ADP Aerodynamic design point
α Angle of attack
A Area
AR Aspect Ratio
APU Auxiliary power unit
BLI Boundary layer ingestion
BPR Bypass Ratio
CG Center of Gravity
CD Coefficient of Drag
CL Coefficient of Lift
ρ Density
d Diameter
e Efficiency factor
F A A Federal Aviation Administration
F AR Federal Aviation Requirements
F SI Fluid Structure Interaction
F Force
g Gravity
h Height, Altitude
HT S High temperature superconductors
h Horizontal
IC AO International Civil Aviation Organization
L/D Lift to Drag Ratio
L Lift
M Mach Number
m Mass flow rate
MT OW Max Takeoff Weight
M AC Mean aerodynamic chord
M f f Mission Fuel Fraction
N LF Natural Laminar Flow
OEW Operating Empty Weight
OPR Overall pressure ratio
S Planform Area
RP M Rotations per minute
a Speed of sound
T ODP Takeoff design point
T O Takeoff
T SFC Thrust Specific Fuel Ratio
T Thrust
t Time
Pt Total pressure
Tt Total temperature
v Vertical
W Weight
Swet Wetted Area
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ContentsPage
1 Introduction 11.1 Summary of Mission Statement, Airplane Design Requirements . . . . . . . . . . . . . . . . . . . 1
2 Concept Evolution & Final Configuration 12.1 Market Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.2 Competing and Baseline Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22.3 Computational Tools . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.4 Sizing Diagrams . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.5 Description of Major Design Decisions and Trade Studies . . . . . . . . . . . . . . . . . . . . . . . 4
2.5.1 Fuselage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.5.2 Propulsion Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52.5.3 Fuel Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
2.6 Tables of Critical Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
3 Propulsion System 63.1 Engine Details . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
3.1.1 Transmission Type Trade Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63.1.2 Powerplant Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
3.2 Fan Blade Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 93.3 Battery and Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9
4 Weight and Balance 104.1 Component Weights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104.2 Center of Gravity Excursion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
5 Aerodynamics 115.1 Airfoil Selection & Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
5.1.1 2D Airfoil Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115.1.2 3D Wing Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12
5.2 High-lift System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135.3 Airplane (3D) Lift Curves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.4 Airplane Drag Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.5 Airplane Drag Polar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145.6 Additional Aerodynamic Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15
6 Stability and Control 156.1 Tail Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156.2 Airplane Trim Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156.3 Static Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166.4 Control Surface Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
7 Performance 177.1 Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177.2 Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177.3 Climb Performance and Cruise Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177.4 Payload-Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
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7.5 Endurance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 187.6 Comparison with Baseline Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
8 Structural Layout, Materials, Manufacturing 198.1 Major Structural Elements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198.2 Material Choices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218.3 Manufacturing Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218.4 V-n Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
9 Systems 219.1 Cabin Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219.2 Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219.3 Emergency Egress . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229.4 Other Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
10 Cost 24
11 Conclusion and Recommendations 24
12 References 26
13 Appendix - Detailed CFD Analysis 2813.1 Surface Pressure Analysis (2D) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
13.1.1 Effect of Sweep . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2813.1.2 Constant AoA Comparison at Cruising Condition . . . . . . . . . . . . . . . . . . . . . . . 2813.1.3 Constant Cl Comparison at Cruising Condition . . . . . . . . . . . . . . . . . . . . . . . . 28
13.2 Flowfield Visualization (2D) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2813.2.1 Flow Contours from Fluent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2813.2.2 Flow Contours from MSES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
13.3 Mesh Quality Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2913.3.1 Convergence Study: Coefficient of Lift . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2913.3.2 Convergence Study: Coefficient of Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2913.3.3 Grids Used (generated via PointWise) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
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List of Figures1 JAG-XP1 Three-View with general dimensions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 Matching diagram. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 Nominal power distribution during cruise. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84 Power distribution during worst case engine out. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95 Carpet plot of possible bypass ratios for varying Mach numbers. Contours represent varying
bypass ratios and inlet Mach numbers. Figure generated with PARA [19]. . . . . . . . . . . . . . 106 Flight Configuration CG Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127 The geometries of two airfoils under consideration are plotted and shown above. . . . . . . . . . 128 The MSES converged 2D Cl/Cd relationship for the HSNLF(1)-0215 and the BACNLF at Re =
30E6 and Ma = .753 are shown above. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129 O-grid mesh with farfield at 50c and y+ initial layer generated via PointWise [9]. Surface pressure
distribution obtained from Fluent [7]. (at Mach .7534, Cp* = -.5796, at Mach .78, Cp* = -.4940) 1310 XP-1 Wing: 3d Lift Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1411 2D and 3D Coefficient of Lift from MSES and Python Iterations . . . . . . . . . . . . . . . . . . 1412 XP-1 Wing: drag curves obtained from by iterating effective angle of attack, circulation, and
coefficient of lift using the Lifting Line Theory. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1513 Trim Diagram for Empty Weight with Static Margin=7.6% . . . . . . . . . . . . . . . . . . . . . 1614 Trim Diagram for Operational Empty Weight and Fuel with Static Margin=10.4% . . . . . . . . 1615 Plot of engine specific thrust as it varies with Mach number . . . . . . . . . . . . . . . . . . . . . 1816 Payload-Range diagram comparing the XP-1 against that of Boeing 757-300 and 757-200 [22]. . 1817 Overall Structure of the XP-1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2018 V-n maneuver diagram at 36,000 f t . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2119 The cabin cross-section of the XP-1 . The elliptical fuselage has the same major-axis dimension
as the widebody Boeing 767 and the same minor-axis dimension as the narrowbody Boeing 737and 757. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
20 The cabin floor plan of the XP-1 . Shown is an all-economy, 28-in pitch configuration whichrepresents the maximum passenger density possible with the aircraft. The seat rows are staggeredto increase elbow room without widening the fuselage. . . . . . . . . . . . . . . . . . . . . . . . . 22
21 Virtual Reality model of internal layout. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2223 The 2014 cost per block hour to operate a Boeing 757-300. Source:planestats.com, Marsh&
McLennan Companies. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2422 The current market price of production aircraft compared to their passenger carrying ability (in
a standard configuration). Up-front competitive pricing for the XP-1, based on this metric, isestimated. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25
24 The sweep angle of 15◦ lowers the Mach number over the 2D section from Mach .78 to .753.The impact of this sweep is shown above. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
25 Surface pressure distribution of the BACNLF airfoil from Fluent (SA and SST) and MSES. . . . 2826 Surface pressure distribution of the BACNLF airfoil using various solvers at identical coefficients
of lift. Note the shock in the MSES solution. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2929 Mesh convergence study of the coefficient of lift using the Spalart-Allmaras turbulence model . . 2930 Mesh convergence study of the coefficient of drag using the Spalart-Allmaras turbulence model . 2931 O-grid-A mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3032 O-grid-A detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3033 O-grid-B mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3034 O-grid-B detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3035 O-grid-C mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3036 O-grid-C detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 30
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37 O-grid-D mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3138 O-grid-D detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3139 O-grid-E mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3140 O-grid-E detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3141 C-grid-1 mesh. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3142 The C-grid-1 detail. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3127 Contours of coefficient of pressure around the BACNLF airfoil during cruise, solved with ANSYS
Fluent using the SA turbulence model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3228 Contours of coefficient of pressure around the BACNLF airfoil during cruise, solved with MSES. 33
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List of Tables1 Mission Specification Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12 Fuel Trade Space . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73 Summary of Aircraft Flight Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74 Summary of Aircraft Lifting Surface Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75 Summary of Powerplant Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76 Summary of Fuselage Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77 Transmission Trade Space . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 98 Powerplant performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99 XP-1 Weight Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1110 Weight and Balance Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1111 XP-1 Parasitic Drag Breakdown at Cruise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1412 Summary of Aircraft Empennage Dimensions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1513 Static Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1714 Takeoff Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1715 Landing Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1716 Comparison with Boeing 757-300 Specifications [22] . . . . . . . . . . . . . . . . . . . . . . . . . 1917 Comparison with Mission Specification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1918 Comparison of Key Performance Parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2019 Energy Consumption of XP-1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2020 Summary of V-n Diagram Design Speeds and Load Factors . . . . . . . . . . . . . . . . . . . . . 2121 XP-1 Cost Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2422 Mesh convergence study raw data using the Spalart-Allmaras turbulence model . . . . . . . . . . 29
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1 Introduction1.1 Summary of Mission Statement,
Airplane Design RequirementsA mission is defined as transporting a certain
payload over a certain distance at a certain cruisecondition between airports of a certain category. Fora passenger transport aircraft, the payload is definedas the passengers, their luggage, and any additionalfreight cargo. The distance is defined as the opera-tional range of the aircraft. The cruise condition isdefined by the cruise speed. The airport category isdefined as the runway length required for takeoff andlanding [1]. A summary of the XP-1 mission is tabu-lated in Table 1. Energy consumption is is defined asthe total amount of energy expended to complete themission, regardless of source.
The baseline aircraft must be a single-aisle trans-port (i.e. Boeing 737) or larger. It must be the mostenergy-efficient aircraft of its size class in service in2005.
Table 1: Mission Specification Summary
SpecificationPayload: 279 Passengers
Crew:1 Flight Attendant per
50 Passengers as per FAA req.Range: 3,500 nmiCruise Altitude: 36,000 ftCruise Speed: M = .78 at 36,000 ft
Take-off:5,000 feet field at an alt. of5,000 feet and a 95 ◦ F day.
Pressurization: 6,000 ft. cabin at 36,000 ft
The XP-1 has the mission goal of transporting279 passengers in order to fill the niche of the Boeing757-300. As per FAA requirements and as stated in 14CFR 91.533, 6 flight attendants are required for 279passengers. Since the design is for 2045, this aircraftwill play a part in transitioning from only human pi-lots to a fully autonomous transport platform. It hastherefore been determined that there will be only onehuman captain pilot (assisted by an AI first officer pi-lot). All humans on board are estimated to weigh 175lbs with 45 lbs of baggage. The total range of 3,500nm was chosen to complete the Boeing 757-300’s In-tercontinental Mission from New York to Anchorage,to Copenhagen, and to Lima, or from Dubai to Hong
Kong, Beijing, and London.It is expected that innovative airframe and
propulsion systems will be implemented and integratedand designed to work in synergy. New operationalparadigms and novel ground infrastructure changes areallowed, but the practicality and implementation chal-lenges of such changes must be analyzed and justified.The design is not required to meet goals for noise orNOx emissions reductions, though these should notbe ignored. The design must consider the needs ofthe customers, development cost and risk, FAA cer-tification, and passenger acceptance. Any design fea-tures are incompatible with current FAA regulations orpassenger expectations must be given special consid-eration. Justification must be provided for the spec-ulations of the advancement and availability of tech-nologies in the 2045 time-frame.
2 Concept Evolution & FinalConfigurationThe design goals for this aircraft are to integrate
the concepts in Strategic Thrust 4 (Transition to Alter-native Propulsion and Energy) with those of StrategicThrust 3 (Ultra-Efficient Commercial Vehicles) of thestrategic implementation plan established by NASA toachieve significant energy efficiency improvements inlong-haul commercial air transport vehicles. The air-craft is to meet NASA’s far-term (beyond 2035) goalfor aircraft fuel/energy consumption, which is 60-80%reduction relative to 2005 best in class aircraft. Theaircraft is also expected to achieve reduction in noiseand NOx emissions.
In pursuing the 60% energy reduction, many ma-turing technologies in machine learning, composites,aerodynamics, and propulsion will be employed. Themomentum of computer advancements over the last20 years have given birth to machine learning, whichhas the capability to evolve into systems such as anAI copilot for the XP-1 . The computer can rapidlychange aircraft settings to adapt to flight conditions.Not only does the aircraft have stall detection capabil-ities, the computer can also deflect control surfaces toalleviate the adverse effects of gusts and adapt the ac-tive cruise flap deflection for current flight conditions.The computer control allows for a smaller static mar-gin, which allows for the implementation of a relativelylow-area V-tail.
The XP-1 is a tri-jet aircraft with two cores and
1
3 fans; the fan is effectively decoupled from the coreelectrically. Recent cryogenic cooling and super con-ductors has allowed to energy conversion with minimallosses. This electric gear of the fan allows for the fan tospin at more optimal speed, and the weight of a tradi-tional gearbox is eliminated. Having three fans allowsfor a large effective bypass ratio, and the third fan onthe rear of the aircraft has no core, therefore effectivelyeliminating the high maintenance costs of a typical tri-jet. With a fan between the V-tail, the take-off noiseis reduced. Since the fans are electric, all three are stillfunctional should one core experience compressor stallor engine surge/backfire. Furthermore, the ability toquickly spool up the fan allows the computer to inducea yaw from the thrust differential, thereby being ableto provide proverse yaw and justify a V-tail, which hasless surface area than a conventional tail but has ad-verse yaw problems. The V-tail is further justified bythe proverse yaw from the wing during a roll; the wingelastically deforms into a Prandtl wing during higher gloadings through advanced carbon fiber tailoring.
The advancements in composites allow for alighter aircraft main structure. The XP-1 has no win-dows and is carbon wound, making the fuselage strongand light. The same increased strength-to-weight ratioallows for a higher aspect ratio wing, which is aeroe-lastically tailored to prevent tip stall, wing divergence,and will exhibit a bell lift distribution during a roll.
An increase in aspect ratio is made possible byadvancements in composites and propulsion. The lowtaper ratio is made possible by aeroelastic tailoring;the low taper ratio lowers weight, a necessity since ahigh aspect ratio generally increase weight. Since thefan is decoupled from the core, the fans under the wingcan be loaded more out board, which also is favorablefor the high aspect ratio.
The wing, fuselage, empennage, and nacelles areall optimized for laminar flow: both natural and active.
2.1 Market StudyThere is an existing gap in the “middle of mar-
ket”, where large narrowbodies and small widebodiescompete to provide airlines with economic solutionsto transport moderate payloads of 220-270 passengersacross moderate ranges of 3000-5000 nmi. This mar-ket segment was previously defined by the Boeing 757(large narrowbody) and 767 (small widebody) families[2].
Currently, Boeing’s 737 MAX-9 and MAX-10
and Airbus’s A321neo and A321LR represent the largenarrowbody competition. They are stretched variantsof the smaller, and aging, 737 and A320 platforms,equipped with modern engines, winglets, cabin interi-ors, and other incremental upgrades [3].
On the other end of the spectrum, Boeing’s 787-8 and Airbus’s A330neo-800 compete in the smallwidebody market. These aircraft are significantlylarger and longer-ranged than the narrowbodies, andshortened versions (e.g. 787-3) never entered produc-tion. The 787 is a new aircraft, featuring mostly car-bon composite construction and other advanced tech-nologies. The A330neo is an re-engined, updated ver-sion of the A330. It is overly expensive to operatethese aircraft on shorter, less dense routes.
There is an operational need for a mid-size air-craft, as large as the 757 and potentially with morerange, to fill the gap in passenger count and range inthe current market. There is a distinct gap in capa-bility between the narrowbodies and widebodies – thisspace is currently serviced by aging 757’s and 767’s orby flying larger aircraft.
2.2 Competing and Baseline Air-craft
The stretched-variant Boeing 757-300 is selectedas the baseline comparison aircraft. Boeing 757 fam-ily production ended in 2004, and no true successoraircraft has emerged since.
The stretched variants of the smaller Boeing 737and Airbus A320 families have come to replace the ag-ing 757’s, but they are still smaller and shorter-ranged.Further stretching is likely not possible without a sig-nificant wing redesign, which Airbus may consider.The largest current narrowbody, the A321LR, is simi-lar to the 757-200 in payload and range, but there isno current match for the 757-300 (or the 767-200).
With a payload of up to 280 passengers in anall-economy, 28-inch pitch layout, the 757-300 has anoperational range of up to 3,235 nmi with a Mach of0.80 and initial cruise altitude of 34,250 ft. This rangeis enough to service trans-Atlantic routes.
Boeing is rumored to be developing the ‘797’,or New Midsize Airplane ‘NMA’, an all-new aircraftwhich can carry 220-270 passengers over a range of5000 nmi. The airplane would replace the 757 and 767families, with a projected entry-into-service in 2025 [4].
Airbus has not yet announced an all-new middleof market aircraft.
2
29'-5"
62'-11"13'-0"
38'-1"
155'-8"47'-10"
12'-9"
59'-9"
145'-0"
16'-6"
31'-11"
13'-2"
140'-8"
Figure 1: JAG-XP1 Three-View with general dimensions.
3
2.3 Computational Tools
A large portion of the analysis was completedusing simple MATLAB tool and many of the diagramsand plots were completed using its built-in functions.
The computational fluid dynamics analysis wascompleted using a wide variety of tools. VSPAEROwas used along with XFOIL to preliminarily analyzethe 3 dimensional lift and drag coefficients of the wing[5]. They were, however, replaced with the MSES tooldeveloped by Mark Drela [6]. Unlike OpenVSP, theMSES tool can perform effective viscous and transonicCFD over our 2D airfoil to more accurately define ouraerodynamic coefficients. The resulting coefficientswere then analyzed in a Python script to obtain 3Dcoefficients. This is further discussed in the Aerody-namics section.
ANSYS Fluent Simulations were run to matchthe MSES output [7]. Fluent runs a RANS solver todetermine the critical flow parameters. Specifically,the settings selected implemented a Strain/Vorticity-Based Spalart-Allmaras turbulence model with aSutherland Law density model. These settings are thestandard for transonic airfoil analysis.
Meshes were generated with Pointwise software,using a y+ wall strategy [8] [9]. Convergence studieswere completed for confidence.
Finally, we run extensive FEA on the airframestructures. We employ the use of ANSYS Structuralwith the ACP Pre and Post processor to analyze ourcomposite structures.
2.4 Sizing Diagrams
The initial weight and sizing study is performedaccording to the methodology set forth by Roskam.The key input parameters are the mission specifi-cation, current technology performance trends, andpredictions of performance improvements with futuretechnology. The primary basis for the analysis is theBreguet range equation, and the goal TSFC, cruiseL/D and Mach. In order to meet the 60% energyconsumption reduction goal, the fuel weight was pre-scribed to be 60% reduced compared to the 757-300. This allowed for significant structural weight de-creases.
The key results are compared with the existingand historic Boeing fleet.
2.5 Description of Major Design De-cisions and Trade Studies
2.5.1 FuselageThe first major design decision to be made is the
selecting the width of the fuselage.A simple parametric study is conducted to ana-
lyze the effect of fuselage width on the length, finenessratio, wetted area, and estimated drag. The fuselageis approximated as a circular cylinder, with diameterequal to the fuselage width and length equal to theoverall aircraft length. At this design stage, no con-sideration is made for the nose or tail cone sections orany eccentricity in the cross-section.
The Boeing 757-300 baseline dimensions can bereduced to two key metrics: total internal volume andcabin floor area. The total internal volume is repre-sented by the volume of the ‘equivalent’ cylinder, andit represents the fuselage ability to fit the volume ofthe payload and systems. The cabin floor area is rep-resented by the area of the rectangular surface definedby the diameter and length of the ‘equivalent’ cylin-der, and it represents the fuselage ability to contain thecabin floor layout. The baseline total internal volumeand cabin floor area are each fixed, and the fuselagewidth is varied. The resulting fuselage length, finenessratio, wetted area, and estimated drag are plotted.The estimated drag is found using the methodologydescribed by Roskam.
The fuselage width is not infinitely variable; in-stead only discrete values are viable for selection be-cause the fuselage must accommodate an integer num-ber of seats and aisles abreast. The width of the seats,spacing between the seats, width of the aisles, andnumber of seats serviced by an aisle are limited byhuman physiology, customer expectations, and (FAA)regulations regarding safety and aircraft exit times.The fuselage widths of existing aircraft are plottedto represent configurations which are known to meetthese limitations.
The XP-1 uses a 16.5 ft wide fuselage, identi-cal in width to the Boeing 767, in order to reducethe fineness ratio compared to the very long Boeing757-300. This results in a wetted area and skin fric-tion drag penalty, but the shorter and wider fuselageconfiguration is advantageous structurally, which willallow for more fuselage weight reduction. Also, theBoeing 757-300’s immense length is serviced by a sin-gle aisle. Moving to a twin-aisle configuration will de-
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40 50 60 70 80 90 100 110 120 130 1400
0.2
0.4
0.6
0.8
1
1.2
Design Point
Figure 2: Matching diagram.
crease loading and unloading times. The XP-1 cancruise at a lower speed, Mach 0.78, and accomplishthe mission in a comparable time.
The fuselage length of 145 ft, longer than re-quired to match the floor space of the 757-300, is se-lected to accommodate the rear fan, which will reducethe usable fuselage area and volume. The fineness ra-tio is approximately 8.8, which is optimal for reducingskin friction drag when accounting for tail sizing.
The fuselage will be constructed of mostlycarbon-composite materials to achieve upwards of 20%weight reductions compared to the 757 and 767, whichuse aluminum construction. This technology and themanufacturing supply chain required to implement ithave already been utilized for the Boeing 787 [10].
The nose cone will reference a Von Karmanogive, and the tail cone shape is yet to be determinedand will be chosen to maximize natural laminar flow.Due the rear fan location, base drag is anticipated tobe effectively zero.
2.5.2 Propulsion DesignThe XP-1 features a ’tri-jet’ set up with a hybrid
electric cycle. Two turbofan engines rest underneaththe two wing sections while a single electric engine isinstalled at the back of the fuselage. Within the tur-bofan, power is transmitted between the core and thefan electrically. Additionally, excess power is generated
and bled off to both the rear engine and auxiliary sys-tems. Electrical connectors, generators, and motors,will make use of high temperature superconductors fa-cilitated by cryogenic refrigerators. A detailed descrip-tion of the powerplant design and a related transmis-sion trade study can be found in the propulsion systemsection.
2.5.3 Fuel Trade Study
The Joukowski Analysis Group analyzed currentand proposed energy sources for the propulsion of theXP-1: Jet-A (standard kerosene jet fuel), syntheticand bio-fuels, liquefied natural gas (LNG), ethanol,and liquid hydrogen (LH2). The design requirementemphasizes that overall energy consumption is to bereduced, regardless of source or ’cleanness’. So, lique-fied natural gas, ethanol, and liquid hydrogen will notbe viable fuels because of their poor per mass or pervolume energy densities. Synthetic and bio-fuels arebeing developed as drop-in replacements, with identi-cal or near-identical performance to Jet-A. XP-1 willachieve the same total energy consumption reductionregardless of if fueled by Jet-A or a drop-in replace-ment.Mass and Volume Efficiency – Kerosene (or itsdrop-in substitutes) as a fuel has a lower energy perunit mass when compared to hydrogen. Hydrogen, onthe other hand, has a lower energy per unit volume.
5
Naturally, we would want to lower the net mass of thefuel of the aircraft, but increasing the required vol-ume will force us to increase the wetted area and dragof the aircraft. Hydrogen is predicted to require 2%more energy on a 3000-nmi mission compared to tra-ditional fuel, and 28% more energy on a 500-nmi mis-sion. Since our aircraft is designed to service routes<3500-nmi, with most missions being much shorterthan 3500nmi, hydrogen would require more energyper average mission [11].Infrastructure Demands – We do not expectwidespread adoption of hydrogen at the scale whichwould prove it to be viable. In order to properly sup-ply a fleet of aircraft with hydrogen, production farmswould have to be constructed at every airport acrossthe globe. This is due to the fact that it is highlyinefficient to transport hydrogen from centralized pro-duction facilities Hydrogen The fuel would have to betransported cryogenically - a technological challenge,or in gas form, which is infeasible due to the volumeof fuel required.Cost – In the 2045 time frame, we expect JET-A to bethe cheaper fuel option. Though fossil fuel resourcesare depleting, transitions to alternative fuels for ve-hicles in other sectors (automotive, maritime) shoulddecrease demand and keep future costs relatively low.This is in comparison to the extremely large amount ofsunk cost associated with constructing hydrogen gen-eration facilities at every airport. Cryogenic-fueled air-craft also have significantly increased refueling times,reducing the utilization of the aircraft and the abilityfor the airline to generate revenue.Emissions – Kerosene has standard emissions, butwith improved efficiency comes decreased emissions.If LH-2 is produced by renewable energy, it is 100%clean, but this goes back to infrastructure needs. Also,the increase in grid energy which is required to producealternative fuels may be offset by non-renewable powergeneration, which would simply "extend the tailpipe"and not meet the overall energy consumption goal.Conclusion – Table 2 summarizes the trade space forthe fuel selection. A score of 1 is weighted the lowest,while a score of 3 is weighted the highest. Naturally,we find that the JET-A option is the most suitableselection for our design. As such, we use its heatingvalues and energy density for this analysis.
2.6 Tables of Critical ParametersA summary of the XP-1 flight, weight, wing, en-
gine, and fuselage parameters are tabulated in Tables3 through 6.
3 Propulsion System3.1 Engine Details
As discussed earlier, the XP-1 ’s propulsion sys-tem will make use of a unique hybrid electic tri-jetsetup. The Joukowski Analysis Group completed ex-tensive design work to size this plant from the groundup, pulling inspiration from various hybrid electric con-cepts that have been proposed. Broadly speaking, theengines follow the same design philosophy as the restof the aircraft, merging exotic choices with maximizedconservative ones. At first glance, from Figure 1, onemay notice the outward engine placement. The place-ment is mostly driven by two factors: the large di-ameter of the engines and the desire to reduce theroot stresses due to lift. The placement trades be-tween excessive dihedral/gull and structural dynamicconsiderations (vibrational concerns and maneuveringissues). Engine diameters were sized to optimize cruiseperformance. Takeoff thrust considerations were lessconcerning given that we can modulate our bypass ra-tio dynamically. By doing this, we trade significantlyreduced takeoff performance for mildly improved cruiseperformance. Given the nature of our long missions,this trade becomes highly favorable.
3.1.1 Transmission Type Trade StudyTwo types of ducted turbofan engines were con-
sidered from a cost and efficiency standpoint: an im-proved version of an N+1 mechanically geared turbo-fan, or the innovative superconducting turboelectricturbofan which would most likely be introduced in theN+2 time-frame. The mechanically geared turbofanemploys a reducing gearbox to allow the fan to op-erate at lower fan pressure ratios (FPRs) while stillretaining a quasi-direct drive. The turboelectric con-cept employs the use of superconducting generatorsand electric motors; instead of mechanically driving thefan, the gas generator core of these engines drives ahigh-efficiency generator or alternator which produceselectricity to be used by the electric motor driving thefan.
6
Table 2: Fuel Trade Space
Design Mass Efficiency Volume Efficiency Infrastructure Demands Cost Emissions NetWeighting 30% 30% 20% 10% 10% 100%Kerosene, JET-A 2 3 3 2 2 2.5Hydrogen 3 1 1 1 3 1.8
Table 3: Summary of Aircraft Flight Information
Parameter ValueMcr ui se .78hcr ui se 36,000 f tFuel Weight 32,120 l bsFuel Volume 4,614 g alMax Payload 60,385 l bsMax Passengers 280 peopleMax Range 3500 nmMTOW 170,500 lbsOEW 78,000 l bs
Table 4: Summary of Aircraft Lifting Surface Param-eters
Parameter ValueWing Reference Area 1636 f t 2
Real Wing Area 1400 f t 2
Wing Aspect Ratio 14.80Root Reference Chord 19.81 f tRoot Exposed Chord 18.64 f tTip Chord 2.57 f tSpan 157.44 f tEmpennage Area 360 f t 2
V-tail Dihedral 38◦
Table 5: Summary of Powerplant Information
Parameter ValueEffective BPRcr ui se 12T SFCcr ui se .45 lb/s
lb f
Takeoff Thrust (Per Engine) 25,600 l b fCruise Thrust (Per Engine) 1,750 lb fFront Engine Diameter 12 f tRear Engine Diameter 10 f t
Turbojets were not considered as they lack theefficiency required for this type of aircraft [12]. Openrotor concepts, though promising high efficiencies(along with future promises of noise reduction), remainlargely unproven at the transonic speeds we intend to
Table 6: Summary of Fuselage Information
Parameter ValueFuselage Length 145 f tFuselage Major Dia 16.50 f tFuselage Minor Dia 13.17 f tFineness Ratio 8.8
operate at (cruise M = .78 and h = 36,000 f t) [13].We choose to hedge our bets and select a ducted fan.Cost and Maintenance – In terms of cost, the elec-trically geared turbofan offers less up front and opera-tional costs due to the relative simplicity of its designwhen compared to the traditional mechanical gearbox[14], [15]. Manufacturing costs are a major drivingfactor; the cost of grinding the mechanical elementsto the precision required in such a gearbox outweighsthe higher cost of the raw materials in an advancedgenerator and motor [16]. In addition, electric gener-ators and motors can be run for extended periods oftime with relatively little maintenance compared to theneeds of a traditional mechanical transmission [15].Efficiency – For the same fan pressure ratio, the elec-trically geared method also offers a potentially moreefficient solution: the extremely high reduction ratiosrequired to run the fan at the extremely low speedsrelative to the core RPM, coupled with the incredi-bly large horsepower associated with the amount ofthrust we require, forces the mechanical gearbox tobe extremely robust, complex, and consequently, in-efficient [15]. Mechanical losses with such a reduc-tion ratio limit the ability of the gearbox to operateat ultra-low speeds [17]. Furthermore, a good amountof efficiency can be justified from the fact that, in theelectric transmission, the core and fan can run at theirown rotational speed during all stages of the mission,increasing the efficiency of the internal gas generatorstages. Moreover, this feature allows the engine to runat varying bypass ratios, allowing us to design thosegeometries for only a single case.Simplicity and Reliability – The electric transmis-
7
sion requires a much greater complexity and suffersfrom higher risk of failure when compared to thegeared turbofan [18]. Modern day metallurgy analysishas the capability to very accurately predict the life-time and wear of complex transmission systems, evenat the scale we would require in this turbofan [16]. Fur-thermore, the mechanical gearing would not require acryogenic supply to run at acceptable efficiencies. Be-cause of these factors, it is our belief that certifica-tion of the mechanical transmission is expected to bemuch more straightforward. Nonetheless, we intend tomitigate the risks of the electric transmission in otheraspects of our design. The dependency on a cryogenicsubsystem can be leveraged with multiple pump andrefrigerator redundancies with acceptable mass penal-ties. In addition, a ’tri-jet’ configuration would adda redundancy which would mitigate the fact that theelectric transmission may be less reliable. We couplethis rationale with the increased flexibility discussed inthe next section.Flexibility and Synergy – The electric propulsionconcept provides unparalleled flexibility in design whencompared to a mechanical transmission propulsion set-up. The ability to completely decouple power gener-ation and consumption of the systems has immensebenefits for other systems. The main benefit is theelectrical transmission’s ability to dynamically respondto engine out cases. In traditional, mechanical trans-mission set ups, an engine out case at the aircraftscale produces constant yaw which must be counter-acted with rudder trim. In an engine out case with anelectric transmission, with a tri-jet setup, power gen-eration could be properly redistributed entirely to therear engine if one of the front engines (or to the frontengines if the rear engine fails), providing sufficientthrust to keep the aircraft in the air. Figures 3 and4 describe the power distribution in both the nominalcase and the engine out case being discussed
Even in normal flight, we can further exploit thisconfiguration by implementing control systems whichallow us to feather the thrust of each engine to activelyproduce yaw moments when needed.
These benefits from the electric transmissionwould allow us to achieve two major objectives: placethe engines more outboard, and use a theoreticallymore efficient tail configuration — the V-tail — withless rudder control surface area. The former would al-low us to justify more ambitious wing configurations,while the latter could drastically decrease the overall
Figure 3: Nominal power distribution during cruise.
parasitic drag of the aircraft.Conclusion – With these trades in mind, we constructthe trade space in Table 7. Due to the nature of themission, we weight pure efficiency higher than all otherfactors. Second comes flexibility, as it can allow us toreach for more optimized designs in other systems, in-creasing the overall efficiency. Weight comes next,followed by the rest of the parameters. A score of 1is weighted the lowest, while a score of 3 is weightedthe highest.
With these considerations, it becomes clear thata tri-jet hybrid electric high bypass turbofan is the mostoptimal choice for our design. Not only does it providethe raw efficiency we need to fulfill the criteria, it alsoallows us to make more aggressive design choices inother aspects of our design.
3.1.2 Powerplant PerformanceThe hybrid electric cycle has the unique capabil-
ity to dynamically adapt the bypass ratio during themission. Since the generator and fans are decoupled,we can increase or decrease the thrust contributions ofeach component to adjust the performance, allowingus to continuously optimize to the mission require-ments. We intend to exploit this ability by reducingour effective bypass ratio at lower altitude, increasingour specific thrust and decreasing the required massflow rate for the same amount of thrust. This flexibil-
8
Table 7: Transmission Trade Space
Design Cost Efficiency Weight Simplicity Reliability Flexibility NetWeighting 10% 30% 15% 5% 10% 25% 100%Turboelectric 2 3 1 2 2 3 2.3Advanced Geared Turbofan 1 2 2 3 3 1 1.7
Figure 4: Power distribution during worst case engineout.
ity frees us from needing to design our inlet diametersfor max takeoff thrust at sea level; we are not con-strained by choking at the inlet during an engine outscenario. We can simply accelerate less air more ef-fectively with the lower effective bypass. This allowsus to be more aggressive with our cruise engine sizing,allowing us to exclusively optimize fan blade geometry,nacelle diameter, etc. for cruise conditions. Figure 5shows the carpet plot outlining the performance of ourpowerplant. This powerplant and its cycle is designedentirely by the Joukowski Analysis Group, and is not acommercial design which is publicly available.
At lower Mach numbers we switch to a regimeof lower bypass ratio in expectation of an engine outsituation. While this sacrifices efficiency, the portionof the mission is fairly short and does not make an un-acceptable impact on our overall efficiency. Once wereach cruise altitude and speeds, we increase our effec-tive bypass ratio and regain our high efficiency. Cal-
culations were completed using standard engine cycleanalyses [12]. Efficiencies were obtained from a futurehybrid electric concept [15]. Table 8 shows some ofthe critical inputs and outputs of the engine analysis.
Table 8: Powerplant performance
Parameter ValueEffective BPRcr ui se 12Effective BPRT O 6OPRcompr essor 80T SFCcr ui se .45 lb/hr
lb f
T SFCT O .41 lb/hrlb f
(F /m)cr ui se 17.3 lb flb/s
(F /m)T O 38.2 lb flb/s
Cruise Thrust (Per Engine) 1,750 lb fTakeoff Thrust (Per Engine) 25,600 l b fFront Engine Diameter 12 f tRear Engine Diameter 10 f tFront Engine Length 15.6 f tRear Engine Length 8 f t
Note that the aggressive overall pressure ratiosreferenced here are a result of potentially significantstrides in compressor cycle development [20].
3.2 Fan Blade InformationThe XP-1 will feature an advanced composite
fan optimized for ultra low fan pressure ratios. Thespecific number of blades and airfoils have not beenselected at this time. Analysis thus far have been con-ducted with the fan blade efficiencies of a GE90. Weintend to optimize the airfoils specifically for the rota-tional velocities at cruise.
3.3 Battery and Fuel SystemAs previously mentioned, the XP-1stores energy
in both the traditional jet fuel and battery power. Bothwill be stored mostly in the wings, with a minoritystored within the lower section of the fuselage. Tocomplete the mission, the aircraft will carry 32,115lbs of fuel along with 2,500 lbs of batteries for rapid
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10 15 20 25 30 35 40
0.25
0.3
0.35
0.4
0.45
0.5
0.55
0.6
0.65
Figure 5: Carpet plot of possible bypass ratios for varying Mach numbers. Contours represent varying bypassratios and inlet Mach numbers. Figure generated with PARA [19].
maneuvering (weight is encompassed within the pow-erplant weight category). The battery stores allowfor approximately 10 seconds of full thrust power attakeoff to allow the engine to rapidly takeoff withoutrequiring the core to spool up to a power balance.In addition, it allows for approximately 60 seconds oflow thrust increase at landing and takeoff for one fan.In other words, it can power a fan to induce a yawmoment for over 60 seconds. This can be used forlow speed maneuvers, allowing the aircraft to com-plete a landing pattern without relying on the V-tail.It can also be used for gust alleviation at cruise. Dur-ing cruise, the batteries are slowly charged by the gasgenerators to replenish the used energy.
4 Weight and Balance4.1 Component Weights
The first step in determining the weight and bal-ance characteristics of the aircraft was to break theoverall aircraft weight into 11 different groups as de-fined by Roskam [21] and as shown in Table 9.
The empty weight consists of aircraft structures,the operational empty weight consists of the emptyweight along with the crew and trapped fuel and oil,and the maximum takeoff weight consists of the op-erational empty weight fully loaded with passengers,
baggage, and fuel.Each component group was broken down into
a percentage of the maximum takeoff weight basedoff of typical commercial aircraft values. For exam-ple, the fuselage typically makes up about 11% of themaximum takeoff weight. However, other design fac-tors can affect this percentage. The fuselage weight isincreased by 8% for a pressurized cabin. [21]
The center of gravity locations for each compo-nent group was chosen based off of generic aircraftcomponent locations as specified and in Mason.
The individual weight components and their lo-cations along the aircraft fuselage are then used tocompute the moment arm from the nose of the air-craft. The center of gravity was then found by addingup all of the moment arms and dividing the sum bythe weight of the aircraft.
4.2 Center of Gravity ExcursionFigure 6 displays the center of gravity locations
for the different weight configurations. It can be seenthat they all fall within the acceptable minimum andmaximum center of gravity range, which is typicallybetween 12% and 32% of the MAC. This informationis also summarized in Table 10, which also displays thedifferent weight configurations for the baseline plane.[22]
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Table 9: XP-1 Weight Breakdown
Group Weights (lbs) Percentage (%) Xcg ( f t ) Moment (lbf-ft)Fuselage 22,832 11 (+8) 65.25 1.3207e6Wing 16,605 8 72.50 9.8820e5Empennage 5,189 2 130.50 4.4469e5Engine 22,832 11 65.25 1.4675e6Landing Gear 7,2645 3.5 75.99 2.9630e5Fixed Equipment 15,775 8 65.25 8.8938e5
Empty Weight (WE):75,778 lbsTrapped Fuel/Oil 1,310 0.77 72.50 9.5001e4Crew 1,075 0.63 8.70 9.3525e3
Operational Empty Weight (WOE):78,000 lbsFuel 32,115 18.85 72.50 2.3283e6Passengers 49,000 28.76 72.50 3.5525e6Baggage 11,200 6.57 72.50 8.1200e5
Max. Takeoff Weight (MTOW):170,380 lbs
Table 10: Weight and Balance Summary
Weight Configuration B757-300 (lbs) XP-1 (lbs) CG Fuselage Location (ft) Percentage MAC(%)Empty Weight 140,000 75,778 73.165 30.7WOE 142,606 78,000 72.969 29.0WOE + Fuel 205,606 110,115 72.832 27.8MTOW - Fuel 210,000 138,265 73.046 29.7MTWO 273,000 170,380 72.943 28.8
The main landing gear of our tricycle configu-ration is located at 55% of the MAC as specified byCurrey’s Aircraft landing gear design. [23] This loca-tion places the landing gear aft of the center of gravityof the aircraft as necessary for stability.
5 Aerodynamics5.1 Airfoil Selection & Characteris-
tics5.1.1 2D Airfoil Selection
Motivated by the reduction in drag from the dragbucket, a high speed laminar airfoil was sought af-ter during the design process. The NASA HSNLF(1)-0215F and NLF(1)-0416 were considered.
The documentation on NLF(1)-0416 indicatethat it is designed for a low Mach and a low Reynoldsnumber, therefore analysis on the airfoil was aban-doned. For both the HSNLF(1)-0215 and the BAC-NLF, the parts of the 2D polars were obtained usingMSES with cruising conditions, and the converged re-sults are plotted in Figure 8. From Figure 8, one can
see that BACNLF has significantly lower 2D drag atcruising conditions; the BACNLF is chosen.
The BACNLF is a Boeing high speed laminarflow airfoil from an European patent (the patent wouldexpire by the time of XP-1). The design condition ofthis airfoil has a coefficient of lift of .50, which is com-parable to that of XP-1 , a coefficient of lift of .43-.48.The root of the wing give a Reynolds number of 28million, which is compatible to the range in which thisairfoil generates large amounts of laminar flow (15-30million). Furthermore, since the XP-1 has a low sweepof 15 degrees, the low t/c of BACNLF (.10) would al-low the wing to have a higher critical Mach numberthan the previously considered NASA NLF airfoils.
According to the patent, the airfoil geometry re-sults natural laminar flow over the upper surface to50% - 60% of the chord in range of Reynold num-ber of 15 million with 25◦ sweep to 30 million with a15◦ sweep [24], and has low losses at the supercriticalMach numbers. On the lower surface, natural lami-nar flow is obtained to about 45% of the chord at thedesign Mach number of .78 and the design CL of .5,
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Figure 6: Flight Configuration CG Locations
0.0 0.2 0.4 0.6 0.8 1.0Non-dim X
−0.2
−0.1
0.0
0.1
0.2
Non
-dim
Y
HSNLF(1)-0213 and BACNLF Geometry
BACNLFHSNLF(1)-0213
Figure 7: The geometries of two airfoils under consid-eration are plotted and shown above.
when the Reynolds number is 30 million [24].The BACNLF-IL airfoil is analyzed Using an O-
grid with y+ initial layer ∆s and a grid control baffleat the shock location to more accurately capture theboundary and shock effects (Appendix, Figure 42). Asshown in Figure 9, the surface pressure distributionis found first with aircraft freestream Mach .78, thenagain with Mach cos(15◦)*.78, or Mach .753. Whilemaintaining the constant Cl , the shock over the topsurface recedes as the Mach number is lowered to ac-count for sweep. This implies that a 15 degree sweepdramatically decreases the wave drag penalties at ourcruising speed. Furthermore, this observation hintsthat a 15 degree sweep is sufficient. From the Cp
plot, the favorable pressure gradient can be observedto the shock location for both cases, so an activelylaminar control may be needed to prevent separation.
Coefficient of Lift (Cl)
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8
Coeffic
ient of T
ota
l D
rag (
Cd)
0
0.005
0.01
0.015
0.02
0.025HSNLF-0215F and BACNLF-IL Drag Polar Comparison
NASA HSNLF-0215F
BACNLF-IL
Ma = .753
Re = 30 x 106
Figure 8: The MSES converged 2D Cl/Cd relationshipfor the HSNLF(1)-0215 and the BACNLF at Re =30E6 and Ma = .753 are shown above.
5.1.2 3D Wing DesignThe XP-1 wing has an aspect ratio of 16, which
is higher than that of all transport aircrafts found to-day. The high aspect ratio wing reduces induced dragand increases the L/D (mainly due to high span), low-ers chord length and Reynolds number, and improvesclimbing performance and decreases airport noise. Thehigh aspect ratio found on the XP-1 is made possibleby a low taper ratio, extended root chords, compos-ite advances, and outboard engine loading (plausiblesince the fan is decoupled from the core, and there isno yaw from the engine surge condition).
The leading edge sweep is lowered to 15 degrees
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x/c
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Cp
-1.5
-1
-0.5
0
0.5
1
1.5
BACNLF 2D Surface Pressure at Mach .78 and Mach .753
Mach = .753 Cl = .473 (Fluent SA)
Mach = .78 Cl = .473 (Fluent SA)
Re = 30 x 106
Figure 9: O-grid mesh with farfield at 50c and y+ ini-tial layer generated via PointWise [9]. Surface pressuredistribution obtained from Fluent [7]. (at Mach .7534,Cp* = -.5796, at Mach .78, Cp* = -.4940)
to reduce crossflow instabilities in an effort to main-tain laminar flow (quarter chord and max thicknesssweep are implied to be even lower). According to a2010 Stanford paper, "wings with low sweep can beeffectively used as an alternative to current sweptbackconfiguration," and it is possible to "significantly re-duce wing sweep without incurring either aerodynamicof structural penalties, especially for M ≤ .8 aircraftdesigns." [25] The same paper cites that the optimalsweep for Mach .81 flight is 15◦, Mach .80 as 10◦, andMach .78 as 5◦ [25]; the XP-1 flies at Mach .78, thereshould be no aerodynamic penalties from the loweredsweep.
The XP-1 employs the gull wing to accommo-date the large fan under the wing and give groundclearance. The resulted dihedral gives stability and de-creases from 12◦ to 5◦ from the root to the mid-spanaileron location.
A low taper ratio of .15 is used to reduce wingweight. Since the taper ratio is low, extra care is putinto making sure the tip does not stall. The tips arewashed-out and aeroelastically tailored for this reason.The twist distribution is set up to provide elliptic liftdistribution during 1g loading. At higher g loadings,the wing is aeroelastically tailored to elastically deformand yield the Prandtl bell lift distribution, which pro-vides reductions in induced drag and also a proverseyaw [26] to assist the V-tail. In designing the twist,XFLR is used to adjust the twist to match the elliptic
distribution. However, since XFLR fails to account forcompressibility, the twist is adjusted with the Prandlt-Glauert correction factor corresponding to the cruisingMach number (this can be done since the correctionaffects the Cl -α slope).
The XP-1 wing is equipped with cruise flaps totake advantage of the AI pilot. The cruise flap will beused to optimized the wing and shift the drag bucketfor cruising conditions, and it will be further discussedin later sections.
Currently, the wing is designed with a constantairfoil, however, for structural and fuel volume con-siderations, the next step is to alter thickness ratiosdistribution and camber of the airfoil while optimizingperformance. The goal is to bump up the thicknessratio at the root to about .13 and lower the thicknessratio at the tip to about .08. Using the Stanford SU2airfoil optimization tool, more optimized airfoils canbe created while maintaining desired thickness ratios.
5.2 High-lift SystemIn order to meet the runway length requirements
for take-off and landing as stipulated by FAR Part 25,the aircraft must have a sufficiently high CLmax . TheCLmax from the airfoil alone was calculated using Open-VSP to be 1.2 at a stall angle of 12◦. The preliminarysizing results, however, resulted in wing-loading andthrust-to-weight ratios that require a CLmaxT O
of 2.2and a CLmaxL
of 2.6. This means that a ∆CLmax of 1.4at landing is required to be able to adequately performlanding and take-off maneuvers.
The potential increases in the maximum lift coef-ficient were calculated separately for landing and take-off, to take into account the variable flap configura-tions. The leading edge features a flap across 80% ofthe span, and the trailing edge has a double-slottedflap across 70% of its span. In order to preserve thelaminar flow, which is a key feature of this aircraft, aKruger flap was seen to be most suitable for the lead-ing edge. Not only does it protect the leading edgefrom becoming rough due to bugs, it has also beenshown that there is potential to provide net gains sim-ilar to leading edge slats when optimized properly. [27]The numerical analysis was conducted for slats, withthe procedure outlined by Scholz, and results in a gainin CLmax of 0.54 at take-off and 0.45 at landing. [28]Using the innovative approach that aims to optimizethe deployment angle and subsequent lift impact, theKruger flap is estimated to improve the CLmax by the
13
same amount.In order to utilize the trailing edge as efficiently
as possible, a double-slotted flap was chosen in an at-tempt to maintain simplicity while achieving the aero-dynamic goals of the aircraft. The double-slotted flapis not as complex as the triple-slotted variety, and willbe easier to package into the relatively slim wing air-foil. Using the flap will further increase the CLmax by0.72 at take-off and 1.0 at landing.
Altogether, this means that the CLmax of 1.2 ofthe clean airfoil will be augmented by 1.28 at take-offand 1.45 at landing. This brings the total CLmax to2.48 at take-off and 2.65 at landing. This satisfies ourneeded values of 2.2 and 2.6.
5.3 Airplane (3D) Lift CurvesFigure 10 shows the 3d lift polar for the XP-1
without the high lift devices during subsonic condi-tions, where the angle of attack is the angle of attackof the fuselage. Data is generated from OpenVSP[29]. The high lift devices offset the curve upwardsby 1.28 during takeoff and 1.45 during landing. Alsonote that OpenVSP does not support stall detection.The aircraft is expected to stall at approximately 12degrees during take off and landing.
-6 -4 -2 0 2 4 6 8 10 12 14
-1
-0.5
0
0.5
1
1.5
2
OpenVSP subsonic 3d lift curve
Figure 10: XP-1 Wing: 3d Lift Curve
Since OpenVSP does not account for compress-ibility effects, an alternative approach to obtaining the3D lift curve was attempted. MSES is used to findthe Cl-alpha curve with cruising Reynolds number andMach number. Unfortunately, the code only convergesbetween -2 and 2 degrees. Using Prandlt lifting linetheory, effective angle of attack, circulation, and coef-
ficient of lift iteration and assuming the 2D Cl curveis the same for all airfoil sections, the 3D Cl is foundby integration through Python. This cruising lift curveover the wing is only valid over a limited range of Angleof Attacks, as seen in Figure 11.
−4 −2 0 2 4Angle of Attack (a)
0.0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
CL
2D and 3D Lift Curves
2D3D
Figure 11: 2D and 3D Coefficient of Lift from MSESand Python Iterations
5.4 Airplane Drag BreakdownTables 11 show the parasitic drag breakdown at
cruise. Data is generated from OpenVSP’s ParasiticDrag analysis tool.
Table 11: XP-1 Parasitic Drag Breakdown at Cruise
Group Cdo (lbs) Percentage (%)Fuselage 0.00327 33.73Wing 0.00232 23.59Front Engines 0.00166 17.14Empennage 0.00106 10.99Rear Engine 0.00056 5.74Engine Pods 0.00018 1.81Total 0.00968 100
5.5 Airplane Drag PolarSince OpenVSP does not account for compress-
ibility, the cruising drag polar will instead by con-structed by using Prandlt lifting line theory and iterat-ing the effective angle of attack, circulation, and coef-ficient of lift. Using the method previously described,
14
the 3D drag polar is found in Python and shown in Fig-ure 12. From this plot, the induced drag coefficient isfound to be around .0040 at cruising CL. Adding thisvalues to the viscous drag coefficient of .0096 fromthe rest of the plane from the cruise drag breakdown,the coefficient of drag (ref. planform area) for the en-tire aircraft is .0141. Dividing the overall coefficientof drag from the overall coefficient of lift, the L/D isfound to be 34.
0.000 0.002 0.004 0.006 0.008 0.010 0.012 0.014CD
0.0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
CL
Wing Drag Components
2DInducedViscousTotal
Figure 12: XP-1 Wing: drag curves obtained fromby iterating effective angle of attack, circulation, andcoefficient of lift using the Lifting Line Theory.
5.6 Additional Aerodynamic Consid-erations
A straight Yehudi is implemented to providestructural support and reduce the interference dragbetween the fuselage and wing root, while prevent-ing fuselage induced separation at the wing root. TheYehudi is made straight to provide maximum flap ef-fectiveness.
The design of the fuselage nose-cone referencesthe Von Karmon Ogive to minimize wave drag. TheXP-1 also uses the fuselage-wing fairing and the under-wing Kuchman bodies to round out the area distri-bution for the transonic area rule, thereby decreasingwave drag. The Kuchman bodies also serve to preventtransonic separation.
The design of the fairing, winglet, and the lami-nar nacelle and pylons will be optimized by future high-fidelity CFD tools.
The integration of the rear nacelle and the tail-cone will depend on the maturity of boundary-layeringesting engine technology.
6 Stability and Control6.1 Tail Sizing
The airplane design in question utilizes a V-Tailfor its favorable geometric, and subsequently aerody-namic, characteristics. The main drawback with a V-Tail was always considered to be its lack of controlla-bility and its negative yaw-roll coupling. With the tri-jet layout of the XP-1 however, thrust vectoring fromthe wing-mounted electric fans can help mitigate suchadverse effects. In addition, the highly sophisticatedcontrol software and pilot aids will allow for safe ma-neuvering of the V-Tail. The unique propulsion layoutalso allows for less vertical tail in the V, as the result-ing moments in an engine-out scenario are not as largedue to the potential support from the engine which isconveniently center-mounted in the rear between thetwo V-Tail surfaces. This means that we have takenan aggressive approach in reducing the vertical portionof the tail by 50%. In addition, the smart controls willallow for an overall reduction of 15% in the overall areaof the tail. This was surface area previously needed tokeep the plane under control, and were one of the rea-sons that the V-Tail advantage was minimized in thepast.
Table 12: Summary of Aircraft Empennage Dimen-sions
Parameter ValueV-Tail Area 410 f t 2
Area of each Surface 205 f t 2
Tail Aspect Ratio 4Tail Dihedral 37◦
Span 29 f tAirfoil thickness 0.09Taper Ratio 0.5
Using conventional tail sizing based off wing-geometry and airplane configuration, the resulting areacould be projected into the necessary V-Tail geometry.Using the reductions previously discussed, the resultingV-Tail has characteristics tabulated in 12.
6.2 Airplane Trim DiagramThe longitudinal stability is analyzed for the
cases of lowest and greatest static margin. From 13,the lowest static margin is 7.6% for the empty weight
15
configuration. The highest static margin is 10.4% forthe operational empty weight and fuel configuration.
−5 0 5 10−0.6
−0.4
−0.2
0
0.2
0.4
0.6
0.8
α in degrees
Cm
Cm
vs α Trim Conditions at Most Aft CG
Elevator Trim −10 degreesElevator Trim −5 degreesElevator Trim 0 degreesElevator Trim 5 degrees
Figure 13: Trim Diagram for Empty Weight withStatic Margin=7.6%
−5 0 5 10−0.8
−0.6
−0.4
−0.2
0
0.2
0.4
0.6
0.8
α in degrees
Cm
Cm
vs α Trim Conditions at Most Forward CG
Elevator Trim −10 degreesElevator Trim −5 degreesElevator Trim 0 degreesElevator Trim 5 degrees
Figure 14: Trim Diagram for Operational EmptyWeight and Fuel with Static Margin=10.4%
6.3 Static MarginTable 13 summarizes the static margin values for
each weight configuration. The static margin is de-fined as the distance between the aerodynamic centerof the entire aircraft and the aircraft center of grav-ity. The aerodynamic center remains constant, butthe center of gravity varies for every weight configu-ration. The aerodynamic center of the aircraft mustalways be aft of the center of gravity to be staticallystable. The aerodynamic center for our aircraft is lo-cated 74.0492 ft from the nose. Comparing this valueto the center of gravity locations in table 10, it is clearthat the aerodynamic center is aft of the center ofgravity for all weight configurations. Therefore, theaircraft is statically stable for all loading conditions.Additionally, though a higher static margin results ingreater longitudinal stability, the static margin of anaircraft should remain between 5-10% of the MAC in
order to maintain a balance between the stability andcontrol. From Table 13, it can be seen that all ofthe weight configuration static margins fall within the5-10% range.
Figure 13 demonstrates trim diagram for theempty weight, and Figure 14 demonstrates the trimdiagram for the operational empty weight with fuel.These two diagrams represent the longitudinal sta-bility for the case of the lowest and highest staticmargin. Both cases have a negative moment coef-ficient vs. angle of attack angle, which means thatthey are both statically stable. The operational emptyweight and fuel case with the higher static margin hasa greater Cm vs α slope than the empty weight casewith the lower static margin. This is consistent withwhat we know about stability because a higher staticmargin results in greater stability, and the steeper sloperepresents a greater tendency for the plane return totrimmed flight when disturbed.
6.4 Control Surface Sizing
To determine the aileron sizing for roll stability,the aircraft is first defined as a Class 3 aircraft. Thismeans that the aircraft is large, heavy low-to-mediummaneuverability aircraft that weighs more than 30,000kg (66,100 lbs). The control surfaces are designed tolevel 1 standards, which require the flying qualities tobe adequate for the mission flight phase. The criticalflight phases are A and C. The XP-1ailerons are de-signed for flight phase C. The roll requirement for aClass 3, Level 1, phase C is 2.5 seconds to achieve a 30degree bank angle at high speeds. [30] The geomet-ric limits to the design are the wing rear spar locationat 75% of the chord and the the outboard flap posi-tion, which is at 70% of the wing span. Therefore, theaileron inboard position is at 70% of the half span oneach side and 95% half span on each side. The aileronchord to wing chord ratio of 0.25, and the correspond-ing aileron effectiveness parameter for this ratio is 0.45.A typical maximum deflection angle for an aileron is30 degrees. [30] +/- 25 degrees is chosen for the XP-1. With these geometric restrictions, the XP-1time toachieve a 30 degree bank roll is 2.2 seconds. This isenough to satisfies the roll requirements.
16
Table 13: Static Margin
Weight Configuration Static Margin (%)Empty Weight 7.6Operational Empty Weight 9.3Operational Empty Weight + Fuel 10.4Maximum Takeoff Weight 9.5Maximum Takeoff Weight - Fuel 8.6
7 Performance7.1 Takeoff
In order to analyze the takeoff performanceof the aircraft, the conceptual approach outlined byRaymer was used. [31] The main assumptions werethat there is no crosswind, and the runway is a dry,hard surface. To make the calculations applicable inthe broadest sense possible, the analysis was done at8,000 ft altitude. It is required by FAR 25 that theaircraft clear a 35 foot obstacle, and this limitation isused to calculate total takeoff distance. Since the air-craft has fairly powerful engines, and is light comparedto its baseline, the distance was well short of the 5,000ft limit. When compared with the takeoff distance of7,700 feet on a standard day at sea level for a Boe-ing 757-300, it is a very small number. At 8,000 feet,the Boeing 757-300 needs 15,500 feet to takeoff. [22]This means that the landing calculation for the XP-1is likely optimistic, but not unrealistic.
Table 14: Takeoff Parameters
Takeoff ParametersWeight (lb) 170,380Rolling Coefficient 0.05Thrust (lbf) 76,800Take-off velocity (ft/s) 239Total distance (ft) 3,300
As is presented in 14, the total takeoff distance is3,300 feet, which satisfies FAR and RFP requirements.
7.2 LandingThe landing analysis was performed with the
same ambient conditions as the takeoff calculations.In this case it is necessary for the aircraft to clear a 50ft obstacle prior to landing on the runway. The theo-retical landing distance is scaled at the end to accountfor the individual variation in piloting style and skill.
The Boeing 757-300 distance for landing at 8,000 feetin altitude is 6,800 feet, and 5,600 at sea level. [22]
Table 15: Landing Parameters
Landing ParametersWeight (lb) 143,000Approach velocity (ft/s) 282Approach Angle (degrees) 3Flare velocity (ft/s) 267Touchdown velocity (ft/s) 250Total distance (ft) 2,600
As is presented in 15, the total landing distance is2,600 feet, which satisfies FAR and RFP requirements.Since both the landing and takeoff distance exceededthe requirements by a fair margin, further concept iter-ations could yield some possibilities for the expansionof some other systems or the further reduction of nec-essary fuel.
7.3 Climb Performance and CruisePerformance
Since most have variable climb performancebased on ambient conditions and airline policy, themaximum rate of climb and climb angle are calcu-lated. This does not mean that the XP-1 will nec-essarily utilize this ability in actual service. The idealclimb angle is over 22 degrees, while the maximum rateof climb associated with this angle is 4,500 feet-per-minute (fpm). It is difficult to find specific numbersfor climb performance of the Boeing 757-300, as thisinformation is largely proprietary. It must also be re-alized, that most commercial jets do not climb at theoptimal rate, and therefore the ideal climb values ofthe XP-1 will be diminished. In terms of climb tra-jectory, the airplane is limited to 250 knots when itsaltitude is under 10,000 feet, with some other localrestrictions at various airports. After that, the pilotis usually instructed to climb as close to the optimum
17
climb velocity as possible (although this value contin-uously changes), until cruise altitude and speed arereached.
0 0.2 0.4 0.6 0.8 1
10
15
20
25
Figure 15: Plot of engine specific thrust as it varieswith Mach number
FAR Part 25 requires that the operational ceilingis set where the airplane is still able to climb at 500fpm. This occurs at 43,000 feet for the XP-1 . As15 shows, this limit is imposed partially by the specificexcess power that the aircraft has at different altitudesand speeds. The calculated ceiling is in a reasonableregime given the cruise altitude of 36,000 feet. Whileoperating at this cruising altitude, the Mach numberwill be 0.78, the L/D of the is 30 and the CL is about0.5.
7.4 Payload-RangeThe Payload-Range characteristics of the XP-1
is compared to that of Boeing 757-300 and 757-200in Figure 16. In calculating the XP-1 range, estimatedfuel used for ground operation, extra loiter, alternate,and descent does not credit to cruising range. As canbe seen from Figure 16, the XP-1 has a longer rangethan both variants of the Boeing 757 in full payloadand full fuel configuration; the XP-1 also has a longerferry range than the 757’s when flying full fuel withno payload. While the XP-1 carries the same maxi-mum number of passengers as the 757-300 (280), theaircraft does not freight as much extra payload.
7.5 EnduranceThe XP-1 has a 7.2 hour cruise endurance in the
full payload maximum take-off weight configuration.
0 1000 2000 3000 4000 5000 6000
Range (nmi)
0
1
2
3
4
5
6
7
Paylo
ad (
lbs)
104 Payload-Range Diagram with Baseline Aircraft
B757-300
B757-200
XP-1
Figure 16: Payload-Range diagram comparing the XP-1 against that of Boeing 757-300 and 757-200 [22].
In the ferry configuration (no payload full fuel), theaircraft has a cruise endurance of 10.5 hours. Onlyfuel used for cruise and climb is used to calculatedrange and cruise endurance, this does not include thefuel estimated for ground operation, loiter, alternate,and descent.
7.6 Comparison with Baseline Air-craft
A quick comparison of the XP-1 and 757-300specifications are shown in Table 16. Note that whilethe XP-1 has a lower max payload weight, the aircrafthas an approximate fuel burn per seat reduction of62%.
All mission specifications are effectively met, thetarget and design mission specs are juxtaposed in Ta-ble 17. The total loading and unloading time savingsfrom the double isle is realized to be 15 minutes; thistime saving in turn allows for a slightly lower cruisingvelocity. As per calculations, a 15 minutes ground sav-ing from a 3,500 nm flight will allow the cruise speedto be lowered to be Mach .78 while maintaining thesave travel time. With the windowless fuselage carbonwinding technology, the cabin can be more aggressivelypressurized to provide passenger comfort.
Table 18 cites some of the key performance pa-rameters of the XP-1 and the Boeing 757-300. Unfor-tunately some information for the baseline aircraft isnot available, as it is either linked to military contractsor simply not released.
The goal of the design is to reduce the energy
18
Table 16: Comparison with Boeing 757-300 Specifications [22]
Parameter 757-300 XP-1Passengers 280 Passengers 279 PassengersICAO Catagory D/4 D/4Cargo Volume 2,370 f t 3 TBDEngine 2x PW2043 Turbo-Electric: 2 cores, 3 fansSL Standard-Day Takeoff Thrust 43,500 lbs 23,867 lbsMax Payload 68,000 lbs 60,200 lbsMax Takeoff Weight 273,000 lbs 170,380lbsOperating Empty Weight 142,060 lbs 78,065 lbsMax Fuel Capacity 11,466 Gallons 5,984 GallonsFull Payload Range 2,650 nm 3,500 nmFull Fuel MTOW Range 3,300 nm 4,765 nmFerry Range 4,250 nm 5,425 nmCruise Mach .8 .78Initial Cruise Altitude 34,300 ft 36,000 ftFuel burn / seat (per 1000 nm) 74.5 lbs 28.1 lbs
Table 17: Comparison with Mission Specification
Mission Goal DesignPayload: 280 Passengers 279 PassengersCrew: 2 Pilots and 6 Flight Attendants 1 Pilot and 6 Flight AttendantsRange: 3,235 nmi 3,500 nmiCruise Altitude: 36,000 ft 36,000 ftCruise Speed: M = .80 at 36,000 ft M = .78 at 36,000 ft
Take-off:FAR 25: 5,000 feet at an altitude of
5,000 feet and a 95◦ F day.FAR 25: 5,000 feet at an altitude of
5,000 feet and a 95◦ F day.Pressurization: 6,000 ft. cabin at 36,000 ft 6,000 ft. cabin at 36,000 ft
consumption of the competing baseline aircraft by atleast 60%. Outlined in Table 19 are the numbers interms of fuel, which is the only source of energy forthe XP-1 and the Boeing 757-300, neither of whichrely significantly on sources of electricity or other pro-pellant on the ground. The values are calculated sim-ply obtaining the total volume of fuel used for bothaircraft, subsequently obtaining a resulting net energyper mission, and comparing it directly to the baseline’senergy consumption. We do this because we completethe exact same mission.
8 Structural Layout, Materi-als, Manufacturing
8.1 Major Structural ElementsThe XP-1’s structure will consist of partially of
the classical airframe structures, and partially of mod-
ern composite materials. It will make use of the previ-ously made configuration design choices to achieve ahighly optimal structure. A preliminary model of thestructure can be seen in Figure 17. The main criticalstructural elements are highlighted in blue, while thecabin floor and back wall are highlighted in green.
The wing will be comprised of the traditional riband spar structure. Each wing will feature a main,reinforced spar at the quarter chord to support themajority of the wing load. A supplementary torsionalspar will be located at approximately 75% of chordto counteract the torsional loads encountered duringflight. This spacing provides substantial room for bothfuel stores and high lift devices, as discussed previously.
The spar cross section will be rectangular tomaximize the bending stiffness of the structure, whilethe ribs cross section will follow the contours of theairfoil in order to define the wing’s shape. Lightening
19
Table 18: Comparison of Key Performance Parameters
Performance Parameter XP-1 Boeing 757-300TSFC: 0.45 lbm/h/lbf Not DisclosedTakeoff Distance: 3,300 ft at 8,000 ft alt. 15,500 ft at 8,000 ft alt.Landing Distance: 2,600 ft at 8,000 ft alt. 6,800 ft at 8,000 ft alt.Cruise Speed: M = .78 at 36,000 ft M = .8 at 36,000 ftMax Climb Rate: 4,500 fpm Not DisclosedOptimum Climb Angle: 22.5 degrees Not Disclosed
Figure 17: Overall Structure of the XP-1
Table 19: Energy Consumption of XP-1
Kerosene Jet Fuel Boeing 757-300 XP-1Volume [Gal]: 11,490 4,750Weight [lbm]: 76,980 32,115Energy [Btu/lb]: 1445 597.7Volume [ f t 3]: 1536 635Reduction [%]: Baseline 62%
holes will also be designed into the rib cross sectionto maximize its stiffness-to-weight ratio. Figure 17details the preliminary structural makeup of the wingand highlights the engine hardpoint (blue rib element)as well as the spar junction inside the fuselage.
Figure 17 also showcases the structural synergywith the V-tail and the rear engine. All three elementswill share a hardpoint for robust load transfer to therest of the structure.
The general cross section of the fuselagewill comprise of periodic bulkheads with structuralstringers. The whole fuselage will be overwrapped witha composite skin discussed in the next section. Thethickness of the bulkheads (distance between the OMLand the interior paneling) will be approximately .5 f t .
The spacing and number of these elements havebeen preliminarily designed based off of drawings ofcommon airframe structures. The specific thicknessand number of composite plies to be applied to each
20
of the 2D elements is yet to be refined. Classicalaerospace structure analysis methods will be appliedas an initial refinement step. Then, advanced com-posite FEA commercial codes will be used to furtherrefine the structure and promote efficient load transfer.
8.2 Material ChoicesThe XP-1 is composed mainly of carbon fiber
composites. These materials are found in the fuselagestructure, wing structure, and skin. The use of carbonfiber reduces the weight of the aircraft while provid-ing high strength. It also improves fuel efficiency andreduces emissions. Additionally, a composite fuselageallows for greater hoop stress limits, and they makeit possible to achieve more comfortable levels of pres-sure and humidity compared to metal structure air-craft. [32]
The landing gear is made of primarily highstrength steel in order to withstand the maximum take-off weight loads and landing loads.
8.3 Manufacturing MethodsA majority of the structures, given their nature
as carbon composite, will be constructed using stan-dard composite layup techniques. The fuselage specif-ically will be constructed by over wrapping carbon fibertow over a mandrel. This method has been well devel-oped for a series of applications, including carbon fiberoverwrapped pressure vessels. The other aerodynamicsurfaces will be manufactured through resin infusionand automated tape laying processes.[32]
8.4 V-n DiagramA V-n diagram was created based on the FAR 25
certification basis. It details the structural and aero-dynamic flight envelope of the XP-1as it cruises at theADP altitude of 36,000 f t . Currently, the maximumvelocity is still being investigated. It is believed that aflutter type failure will most likely drive that parameter(due to the nature of the high AR wing), but extensiveFSI analysis is required to properly define it.
Table 20 lists the design speed capabilities andmaximum load factors of the aircraft.
9 Systems9.1 Cabin Layout
The XP-1 ’s elliptical fuselage cross-section em-phasizes passenger-carrying ability instead of cargovolume, since the area which is greater than stand-
0 100 200 300 400 500 600
-2
-1
0
1
2
3
4
Figure 18: V-n maneuver diagram at 36,000 f t
Table 20: Summary of V-n Diagram Design Speedsand Load Factors
Flight Condition Flight Speed (KEAS)Positive 1-g Stall Speed 160 knotsNegative 1-g Stall Speed 226 knotsDesign Maneuvering Speed 277 knotsDesign Cruising Speed 256 knotsDesign Diving Speed 474 knots
Positive Limit Load Factor 3.00Ultimate Load Factor 3.60
Negative Limit Load Factor -1.00
ing height is reduced compared to a circular sec-tion. Carbon-composite construction will alleviate is-sues with the cyclic loading due to the pressurizationof a non-cylindrical vessel during climb and descent.The reduction from a 16.5-ft circular section to a 16.5x 13.2 ft elliptical section will reduce skin friction byan amount roughly proportional to the reduction inperimeter, nearly 10%.
Virtual reality models were also created in or-der to get a better sense of the seat dimensions andpassenger comfort.
9.2 Landing GearThe landing gear will consist of a retractable tri-
cycle configuration with a nose landing gear and mainlanding gear of the same class a Boeing 737-800. Themaximum takeoff weight of the 737-800 is comparableto the XP-1 . The location of the center of gravity of
21
TOP OF ROLLERSTO CEILING
186in
13'2"
28in
16'6"
62in19in
64in
83in
158in
198in
44in18in
42in
EXTERIOR
INTERIOR TRIM TO TRIM15'6"
33in
67in
LOOKING AFTECONOMY CLASS SEATING
SEVEN-ABREAST
Figure 19: The cabin cross-section of the XP-1 . Theelliptical fuselage has the same major-axis dimensionas the widebody Boeing 767 and the same minor-axisdimension as the narrowbody Boeing 737 and 757.
the main gear along the fuselage length was chosenat 55% of the mean aerodynamic chord as suggestedin Currey’s Aircraft Landing Gear Design. [23] Thislocation places the landing gear 2.8 feet aft of theupper center of gravity limit and 3.1 feet aft of thelower center of gravity limit. The height of the land-ing gear is designed to satisfy the FAR Part 25.925requirement that states that the propeller, or in thiscase the engine nacelle, needs to have a 7 inch clear-ance from the ground. However, a ground clearanceof 3 feet is used for passenger safety. This results ina landing gear height of 15.1 feet from the center ofthe fuselage. The tipback angle for this height is 10.6degrees, which is greater than the maximum takeoffangle of 10 degrees. Additionally, the takeoff rotationclearance angle is 14.7 degrees, which is also greaterthan the 10 degree takeoff angle.
The nose gear is required to support greater thanor equal to 5% and less than or equal to 20% of theaircraft weight. If the nose supports 5% of the weight,the base wheel base length (distance between mainlanding gear and nose landing gear) is 62.9 feet. Thisbase length is chosen to maintain the minimum andmaximum CG of the aircraft within the CG limits.
The track length is designed so that the lateraldistance between each main gear and the aircraft cen-ter of gravity is greater than 25 degrees. A value of30 degrees is chosen to be safe. The track length to
LG
L
LL
A
A
A
AG
C
C
OVER-WINGEMERGENCY EXIT
L
DOOR
70'0"
121'4"
145'0"
19'10"
G - GALLEYL - LAVATORY
A - ATTENDANTC - CLOSET
INTERIOR ARRANGEMENTS - ALL-ECONOMY CLASS CONFIGURATION279 SEATS AT 28-IN PITCH
Figure 20: The cabin floor plan of the XP-1 . Shownis an all-economy, 28-in pitch configuration which rep-resents the maximum passenger density possible withthe aircraft. The seat rows are staggered to increaseelbow room without widening the fuselage.
Figure 21: Virtual Reality model of internal layout.
satisfy this has to be equal to or greater than 14.7feet. [30]
9.3 Emergency EgressAs shown in Figure 20, there are four main doors
(two Type A passenger doors and two Type A servicedoors) and two overwing emergency exit doors (TypeIII). As such, 48 inch and 30 inch provisions, respec-tively, have been made in the floorplan. The configura-tion is similar to the Boeing 767-200 and is expected tomeet FAR 25.8xx regulations and be compatible withairport boarding gangways.
9.4 Other SystemsThe XP-1 will feature numerous advanced sys-
tems which both improve synergy and efficiency.
22
Active Laminar Flow – The XP-1 aircraft will beequipped with laminar flow control devices. As seenfrom the 757 HLNLF flight test from 1990, the suctiondevices at the leading edge can delay transition to over65% of the chord at the design condition of CL = .5and Mach .80, a speed comparable to that of XP-1.[33] Suction panels will be placed under the leadingedge of the the empennage to maximize laminar flow.Such laminar flow control on the fuselage can also give30% laminar flow over the fuselage on business jets,at the operating conditions of the XP-1 a higher per-centage of laminar flow is anticipated. [34] Accordingto a newly published article by TU Braunschweig, suc-tion from 7% to 70% of the fuselage length can delaytransition to 80% of the fuselage length as per CFDprediction.
Laminar flow over the nacelles will be obtainedfrom geometric optimization from faster high-fidelityCFD softwares made available in the coming years;cryogenic cooling of the engine core nacelles will alsobe employed to maximize laminar flow.Battery & APU – The XP-1 will not require an aux-iliary power unit (APU). The two hybrid electric tur-bofan cores at the front, which already have the abilityto produce excess energy for the rear fan, will be de-signed such that the optimal load on them includesthe power draw from both the propulsion and auxil-iary power systems. The rationalization to not havean APU allows us to use the saved weight on a mod-erately sized battery supply. This supply allows thefans to instantaneously draw power as needed with-out requiring the turbofan cores to slowly spool upto meet the power balance needs. This enables thefans to rapidly throttle up when quick bursts of thrustare needed, such as during takeoff, balked landing, orwhen rudder moments are needed (accomplished bythe aforementioned purposeful asymmetric thrust).Cryogenic Cooling – Cryogenics will be used usedthroughout the aircraft to take advantage of both theactive laminar flow and superconducting technologiesdiscussed previously. The addition of this system willcome with an approximately 8,000 lb weight penaltyalong with an overall .4% decrease in overall propul-sive efficiency due to the increased power draw [15].All-electric Secondary Systems – No bleed off airfrom the compressor stage will be used for secondarysystems. All secondary systems which traditionallymade use of this technology will be powered electri-cally. We estimate that this shift allows us to justify
at least 2,000 lbs of empty weight reduction in addi-tion to gaining some efficiency within the powerplant[35].De-icing – The use of cryogenics for the ALF andsuperconducting engine systems puts the nacelle andwing surfaces at high risk for icing. Ice formations onthese surfaces could trigger the turbulence early alongthe length, negating the benefit of the technology. Inaddition, icing off of the top of the fuselage puts therear engine at risk of ice ingestion As such, we intendto implement active ultrasonic anti-icing technologiesalong the critical surfaces. These devices will applyhigh frequency vibrations to the at-risk surfaces andinstantaneously de-ice them when the ice formationson them become large enough to effect the aerody-namics or pose a threat to the rear engine [36]. Powerconsumption of this system is expected to be minimal:.5 W /i n2 [36].Cruise Flaps – Since the XP-1 utilizes a natural lam-inar flow airfoil in order to minimize drag, there is apenalty in the low-drag regime across varying lift coef-ficients. The drag-bucket can be widened, however,with the use of cruise flaps at the trailing edge ofthe wing. [37] Automatic cruise flaps, like the onesproposed by Gopalarathnam and McAvoy, can adjustthe airfoil in such a manner that results in a favor-able pressure gradient on both the upper and lowersurface. The cruise flap aims to move the stagnationpoint to a small area at the leading edge where thiscondition is desirable. In order to automate the deflec-tion of the cruise flaps, a control system needs to beimplemented in order to avoid the high pilot workloadassociated with manual deflection. This system can bebased on various input parameters that represent theambient conditions, with top and bottom surface pres-sure being proposed by Gopalarathnam and McAvoy.[37] Structural considerations must also be made whenconstructing the existing trailing edge flaps in orderto incorporate cruise flap functionality. The aft flapmay need to deflect by itself, or contain a separatesection that can be controlled independently duringcruise, where the normal flaps are not needed. The re-sult of this effort will be that the aircraft can cruise atlow Cd across a wider spectrum of Cl and thus acrossmore speeds.
23
10 CostIt is uncertain what the XP-1 will cost, but it
must not cost excessively more than the competingaircraft or else airlines will not purchase it. The XP-1will offer reduced fuel costs per block hour or avail-able seat mile (ASM) by 60%. These savings ac-cumulate over the lifespan of the aircraft (estimatedas 20 years). This will justify an increased up-frontcost to the airline and a higher manufacturing cost forthe airframe producer, but with increased up-front in-vestment comes increased interest. The Boeing year2036 market outlook anticipates 32,190 narrowbodyjets and 5,710 small widebody jets [38]. The XP-1market is estimated to include the largest 10% of nar-rowbody aircraft and the smaller 33% of small wide-body aircraft - a total of ≈ 5000 jet sales. The R&Dcost of the airplane program and the per-aircraft acqui-sition cost are estimated using the method outlined inAirplane Design: Part VIII by Roskam. The method isa strong function of historic aircraft weight and speedtrends as well as an assumed technological advance-ment factor and material cost factor. Cost estimateA follows the Roskam methodology exactly, and re-sults in a low estimate, largely due to the low weightof the XP-1. Cost estimate B uses a doubled tech-nological advancement factor and material cost factorto account for the high technology level. Finally, costestimate C assumes an aircraft weight equal to theBoeing 757-300. These corrections still result in sur-prisingly low estimates for RDTE and per-unit cost.The Boeing New Midsize Airplane is estimated by in-dustry analysts to have over a $10B development cost,and the airplane should cost appreciably more than thesmaller, less capable $130M Boeing 737 Max10.
Table 21: XP-1 Cost Breakdown
Cost A B CRDTE Cost 3.3B 6.2B 8.2BUnit Price 95M 111M 121M
The per-block hour cost is estimated by assum-ing similar costs to the baseline aircraft. In 2014, Witha per block hour fuel cost of $3818 and a 12.5 blockhour per day utilization, a United Airlines-operated757-300 consumed $17.4 million in fuel. The XP-1would save over $200 million in lifetime fuel expensesin comparison, which would likely pay for the airplane.
Figure 23: The 2014 cost per block hour to operatea Boeing 757-300. Source:planestats.com, Marsh&McLennan Companies.
It is not expected that, except for fuel costs, theXP-1 will have different operational costs compared toexisting aircraft since it uses standard fuel and airportoperating procedures and has a similar crew comple-ment.
11 Conclusion and Recom-mendations
The XP-1 is set to meet the requirements setby the design goals without resorting to extraordinarilyrisky designs. It will be poised to be the prime replace-ment for the 757-300 in the 2038-2045 timeframe forwhich it is designed. It incorporates multiple advancedtechnologies and emphasizes synergy between systemsrather than maximizing gains within singular systems.We do not compromise safety in our designs; rather,it is our believe that this aircraft is more robust inmany aspects, including resistance towards engine outcases and pilot error. While advanced systems oftencome with increased costs, there are careful consid-erations in place that will help to mitigate these. Inaddition to citations from a variety of reputable re-search institutions, all design decisions are backed withstandard calculations. There were no specific hyper-complex systems that required super advanced anal-ysis techniques for this design. When compared tothe baseline aircraft Boeing 757-300, the XP-1 looksto be a prime replacement for the middle-of-marketairline industry while advancing the market in an envi-ronmentally friendly future and adhering to FAR Part25 airworthiness standards.
Additional work is needed in many places. For
24
100 150 200 250 300 350 400 450 500 550 600$50
$100
$150
$200
$250
$300
$350
$400
$450
$500
Figure 22: The current market price of production aircraft compared to their passenger carrying ability (in astandard configuration). Up-front competitive pricing for the XP-1, based on this metric, is estimated.
one, the structures of the aircraft need to be more re-fined, with dimensions and specific material makeupsof all the of the critical elements. For aerodynamics,the 3D lift curves must be refined and better definedfor each of our conditions. Currently, there is only apreliminary analysis for 3D lift curves during landing,takeoff, and cruise. In addition, it may be necessaryto complete a more robust 3D drag analysis by pool-ing computational resources and running 3D FluentRANS simulations. Moreover, an MSES simulation ofthe high lift devices needs to be performed and in-vestigated in order to more closely determine how tocombine the high lift devices and cruise flaps. Despitethese challenges, the careful consideration of all of thevarious design choices and trade-offs result in an air-craft that complies not only with the design challengeset forth by NASA, it presents a legitimate possibilityfor the future of sustainable aviation.
25
12 ReferencesReferences[1] E. Obert, “Aerodynamic design of transport aircraft,” in. Amsterdam: IOS Press BV, 2009, ch. 1,6.
[2] D. Arkell, With the 7e7, commercial airplanes looks to solve the riddle in the middle, 2003.
[3] B. Zhang, Boeing’s newest plane is also its biggest headache, 2017.
[4] C. S. V Bhaskara, Boeing reveals additional details on its possible new middle of the market aircraft, 2017.
[5] M. Drela, Xfoil: An analysis and design system for low reynolds number airfoils, 1989.
[6] ——, Mses: An analysis and design system for single- or multi-element airfoils, 2007.
[7] ANSYS, Fluent 19.0, 2018.
[8] S. M. Salim and S. Cheah, Wall y+ strategy for dealing with wall-bounded turbulent flows, 2009.
[9] Pointwise, Pointwise 18.1, 2018.
[10] D. Watts, P. Sumida, B. Bunin, G. Janicki, J. Walker, and B. Fox, A study of the utilization of advancedcomposites in fuselage structures of commercial aircraft, 1985.
[11] R. Walther, R. Hendricks, D. Daggett, and O. Hadaller, Alternative fuels and their potential impact onaviation - nasa/tm-2006-214365, 2006.
[12] J. D. Mattingly, Elements of propulsion, 2006.
[13] M. D. Guynn, J. J. Berton, M. T. Tong, W. J. Haller, and D. R. Thurman, Initial assessment of openrotor propulsion applied to an advanced single-aisle aircraft, Langley Research Center, 2011.
[14] R. Guha, Reliable bearing wear detection system for on-condition maintenance of electric generators, 2002.doi: 10.4271.
[15] J. L. Felder, H. D. Kim, and G. V. Brown, Turboelectric distributed propulsion engine cycle analysis forhybrid-wing-body aircraft, Glenn Research Center, 2007.
[16] O. d. Narek Shougarian Jeremy Agte, Multistate analysis and optimization of a geared turbofan enginelubrication system, 2012.
[17] G. V. Brown, A. F. Kascak, and B. Ebihara, High power density motors for aeropropulsion, Glenn ResearchCenter, 2005.
[18] R. D. H. T A Lyon, Geared fan engine systems—their advantages and potential reliability, 1973.
[19] J. E. Watkins, Parametric cycle analysis matlab program v1.00, 2012.
[20] A. L. Sascha Kaiser Stefan Donnerhack, A composite cycle engine concept with hecto-pressure ratio,2015.
[21] J. Roskam, “Aiplane design part 2,” in. Roskam Aviation and Engineering Corporation, 1985.
[22] Boeing, Boeing 757 information, 2009.
[23] N. Currey, “Aircraft landing gear design: Principles and practices,” in. AIAA, 1988.
[24] T. C. Versteegh, Natural laminar flow, low wave drag airfoil, 1982.
[25] S. S. Anthony Jameson John C. Vassberg, Aerodynamic–structural design studies of low-sweep transonicwings, Stanford University, 2010.
[26] O. J. M. Albion H. Bowers, On wings of the minimum induced drag: Spanload implications for aircraftand birds, Armstrong Flight Research Center, 2016.
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[27] H. Akaydin, J. Housman, C. Kiris, C. Bahr, and F. Hutcheson, Computational design of a krueger flaptargeting conventional slat aerodynamics, 2016.
[28] D. Scholz, High lift systems and maximum lift coefficients, 2016.
[29] o. s. NASA, Openvsp 13.6.0, 2018.
[30] M. Sadrey, Aircraft design: A systems engineering approach, 2013.
[31] D. Raymer, “Aircraft design: A conceptual approach,” in, J. Przemieniecki, Ed., ser. AIAA EducationSeries. New York: AIAA, 1989, pp. 5, 153.
[32] K. B. Katnam, “Bonded repair of composite aircraft structures: A review of scientific challenges andopportunities,” in Progress in Aerospace Sciences, vol. 61, Elsevier, 2013, pp. 26–42.
[33] R. D. Joslin, Overview of laminar flow control, 1998.
[34] N. Beck, T. Landa, A. Seitz, L. Boermans, Y. Liu, and R. Radespiel, Drag reduction by laminar flowcontrol, 2018.
[35] R. R. F. W E Murray L J Feiner, Evaluation of all-electric secondary power for transport aircraft, 1992.
[36] J. R. Jose Palacios Edward Smith, Instantaneous de-icing of freezer ice via ultrasonic actuation, 2011.doi: 10.2514.
[37] C. McAvoy and A. Gopalarathnam, Automated cruise flap for airfoil drag reduction over a large lift range,2002.
[38] M. A. Boeing Commercial Airplanes, Current market outlook 2017-2036, 2017.
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13 Appendix - Detailed CFDAnalysis
13.1 Surface Pressure Analysis (2D)13.1.1 Effect of Sweep
The BACNLF-IL airfoil is analyzed Using the O-grid as shown in Figure 40 (50c farfield, y+ initialspacing with 1.2 growth factor, and extra nodes atthe shock location). The surface pressure distributionis found first with aircraft freestream Mach .78, thenagain with Mach cos(15◦)*.78, or Mach .753. Usingthe freestream Mach numbers and the specific heatratio, the critical Cp* is -.5796 for Mach .73, and -.4940 for Mach .78.
x/c
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Cp
-1.5
-1
-0.5
0
0.5
1
1.5
BACNLF 2D Surface Pressure at Mach .78 and Mach .753
Mach = .753 Cl = .473 (Fluent SA)
Mach = .78 Cl = .473 (Fluent SA)
Re = 30 x 106
Figure 24: The sweep angle of 15◦ lowers the Machnumber over the 2D section from Mach .78 to .753.The impact of this sweep is shown above.
While maintaining the same Cl, the shock overthe top surface can be observed to recede. This im-plies that a 15 degree sweep dramatically decreases thewave drag penalties at our cruising speed. Further-more, this observation hints that a 15 degree sweepmay be sufficient for the aircraft, however, further 3Danalysis is required to reach a definite conclusion.
13.1.2 Constant AoA Comparison at CruisingCondition
Using the same grid as the previous analysis,the surface pressure is found using both the Spalart-Allmaras (SA) and Menter’s Shear Stress Transport(SST) turbulence models at Mach .753, where theangle of attack is .93◦ (interpolation for target Cl =.48). This is of interest because the while the popular
SA model is more robust and computationally cheapersince it only solves one equation, the SST model ismore physical, albeit less stable in the presence of sep-aration. The pressure distribution from the two turbu-lence models are obtained and compared in Figure 25:the differences between the pressure distributions arefound to be minimal.
Using the same angle of attack, the surface pres-sure distribution is obtained from MSES using thesame Reynolds number and Mach number. Comparedto the distribution found from Fluent, the MSES dis-tribution over predicts the lift and therefore results alarger shock. This discrepancy can be due to farfieldboundary conditions, mesh quality, convergence crite-ria, etc.
x/c
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Cp
-1
-0.5
0
0.5
1
1.5
BACNLF 2D Surface Pressure
MSES (AoA = .94 deg)
Fluent SA (AoA = .94 deg)
Fluent SST (AoA = .94 deg)
Figure 25: Surface pressure distribution of the BAC-NLF airfoil from Fluent (SA and SST) and MSES.
13.1.3 Constant Cl Comparison at CruisingCondition
Using MSES, another surface pressure distribu-tion is obtained to match the coefficient of lift. In Fig-ure 26, the new surface pressure distribution is seen tomatch the distribution obtained from fluent, all but atthe shock location. The MSES pressure distributioncontinues to display a larger shock.
13.2 Flowfield Visualization (2D)13.2.1 Flow Contours from Fluent
Using Fluent, the flow pressure contour is ob-tained using SA at cruising conditions. The contour isshown in Figure 27.
28
x/c
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1
Cp
-1
-0.5
0
0.5
1
1.5
BACNLF 2D Surface Pressure
MSES (Cl = .473)
Fluent SA (Cl = .473)
Fluent SST (Cl = .473)
Figure 26: Surface pressure distribution of the BAC-NLF airfoil using various solvers at identical coeffi-cients of lift. Note the shock in the MSES solution.
13.2.2 Flow Contours from MSESUsing MSES, the flow pressure contour is ob-
tained at cruising conditions. The contour is shown inFigure 28.
13.3 Mesh Quality StudyUsing the O-grids shown in Figures 31 through
40, the coefficient of drag and lift are calculated foreach grid (Ma = .754, AoA = .93, Re = 30E6, SA),and tabulated in the Table 22 respectively to the num-ber of cells.
Table 22: Mesh convergence study raw data using theSpalart-Allmaras turbulence model
Number of Cells Cl CdO-grid A 5,456 .4179 .00903O-grid B 11,808 .4616 .00804O-grid C 25,792 .4608 .00774O-grid D 55,776 .4607 .00746O-grid E 72,964 .4733 .00735
13.3.1 Convergence Study: Coefficient ofLift
In Figure 29, the obtained Cl is plotted over thenumber of elements. The Cl is seen to stabilize untilthe last grid, which was created specifically to captureboundary layer and shock wave effects with greaterresolution.
Number of Mesh Cells ×104
0 1 2 3 4 5 6 7 8
Co
eff
icie
nt
of
Lift
(Cl)
0.41
0.42
0.43
0.44
0.45
0.46
0.47
0.48Convergence Study: 2D Cl
60pps-ds.004
120pps-ds.002
240pps-ds.001
480pps-ds.0005
480pps-60ptShock-ds.00005(y+)
Figure 29: Mesh convergence study of the coefficientof lift using the Spalart-Allmaras turbulence model
13.3.2 Convergence Study: Coefficient ofDrag
In Figure 30, the obtained Cd is plotted over thenumber of elements. While not fully converged, theconvergence plot hints the converged Cd to be some-where near Cd = .0072. Moreover, since the conver-gence is seen to be from high to low, the O-grid E canbe considered to be sufficient and conservative sincefurther resolution increase will only decrease the coef-ficient of drag.
Number of Mesh Cells ×104
0 1 2 3 4 5 6 7 8
Co
eff
icie
nt
of
Dra
g (
Cd
)
×10-3
7.2
7.4
7.6
7.8
8
8.2
8.4
8.6
8.8
9
9.2Convergence Study: 2D Cd
60pps-ds.004
120pps-ds.002
240pps-ds.001
480pps-ds.0005
480pps-60ptShock-ds.00005(y+)
Figure 30: Mesh convergence study of the coefficientof drag using the Spalart-Allmaras turbulence model
29
13.3.3 Grids Used (generated via PointWise)
Figure 31: O-grid-A mesh.
Figure 32: O-grid-A detail.
Figure 33: O-grid-B mesh.
Figure 34: O-grid-B detail.
Figure 35: O-grid-C mesh.
Figure 36: O-grid-C detail.
30
Figure 37: O-grid-D mesh.
Figure 38: O-grid-D detail.
The O-grid E mesh was created to better capturethe boundary layer and shock effects. A baffle wascreated at the shock location to add extra nodes toimprove resolution at the shock location. Using they+ theory [5], the initial ∆s spacing is calculated toaccurately capture wall-bounded turbulent flow.
Figure 39: O-grid-E mesh.
Figure 40: O-grid-E detail.
The t-rex grid shown in Figure 41, will seed theanalysis for airfoil optimization with the SU2 software.
Figure 41: C-grid-1 mesh.
Figure 42: The C-grid-1 detail.
31
Figure 27: Contours of coefficient of pressure around the BACNLF airfoil during cruise, solved with ANSYSFluent using the SA turbulence model
32
Figure 28: Contours of coefficient of pressure around the BACNLF airfoil during cruise, solved with MSES.
33
