Vendetta Final Proposal Part 1 (3.4 MB) - Cal Poly

Figure 11.1 - 1-g Military Specific Excess Power Envelope at Maneuver Weight ............................................................59Figure 11.2 - 1-g Maximum Specific Excess Power Envelope at Maneuver Weight.........................................................60Figure 11.3 - 2-g Maximum Specific Excess Power Envelope at Maneuver Weight.........................................................60Figure 11.4 - 5-g Maximum Specific Excess Power Envelope at Maneuver Weight.........................................................61Figure 11.5 - Maximum Sustained Load Factor Envelope at Maneuver Weight................................................................61Figure 11.6 - Maneuverability Diagram at 15,000 ft and Maneuver Weight......................................................................62Figure 11.7 - Maneuverability Diagram at Sea-Level and Maneuver Weight....................................................................62Figure 11.8 - Fuel Consumption Envelope at Average Climb Weight ...............................................................................63Figure 11.9 - Fuel Consumption over Mission ...................................................................................................................64Figure 11.10 - Takeoff Profile............................................................................................................................................65Figure 11.11 - Landing Profile ...........................................................................................................................................65Figure 12.1 - L to R configurations 1, 2, 3 .........................................................................................................................67Figure 12.2 - 180 inch MPRL.............................................................................................................................................67Figure 12.3 - Ballute and Sabot ..........................................................................................................................................67Figure 12.4 - 30in Ejector Rack..........................................................................................................................................68Figure 12.5 - LAU-142A Ejection Sequence......................................................................................................................68Figure 12.6 - MPRL with 8 × 2,000 lb JDAMs..................................................................................................................69Figure 12.7 – MPRL with 8 × AGM-158A (JASSM) ........................................................................................................69Figure 13.1 - Cockpit Width Trade Study ..........................................................................................................................71Figure 13.2 – Fuselage Comparison ...................................................................................................................................71Figure 13.3 - Virtual Cockpit Model ..................................................................................................................................72Figure 13.4 - Rectilinear Vision Plot of Forward Cockpit Position....................................................................................72Figure 13.5 - Cockpit Displays...........................................................................................................................................72Figure 13.6 - Advanced Fighter Crew Protection System ..................................................................................................73Figure 13.7 - K-36D Performance Envelope ......................................................................................................................73Figure 14.1 - APU Placement.............................................................................................................................................74Figure 15.1 - Routing Tunnel .............................................................................................................................................78Figure 15.2 - Manufacturing Breaks...................................................................................................................................78Figure 15.3 - Assembly Line ..............................................................................................................................................79Figure 16.1 - Cost Analysis ................................................................................................................................................80Figure 16.2 - Operating Cost ..............................................................................................................................................81Figure 16.3 - Lifecycle Cost ...............................................................................................................................................81vi

List of TablesTable 1.I - Required Weapons Loadout................................................................................................................................1Table 1.II - Summary of Design Requirements....................................................................................................................2Table 1.III - Comparison of the F-111, F-117, B-2, B-1B, and F-15E .................................................................................4Table 2.I - Weight Fractions & Weights...............................................................................................................................5Table 2.II - Weight Fraction Assumptions ...........................................................................................................................6Table 2.III - Constraint Assumptions ...................................................................................................................................6Table 2.IV - Initial and Current Sizing.................................................................................................................................7Table 5.I - Wing Measurements .........................................................................................................................................19Table 5.II - Parasite Drag Component Buildup (50,000 ft, Mach 0.5) ...............................................................................25Table 6.I - Engine Specifications of RFP Supplied Engine................................................................................................26Table 6.II - RFP Dimensions Compared to the SNECMA Olympus .................................................................................27Table 6.III - IHPTET Goals................................................................................................................................................28Table 9.I - Initial Component Weight Buildup...................................................................................................................42Table 9.II - Final Component Weight Buildup...................................................................................................................43Table 9.III - Inertia Estimation ...........................................................................................................................................44Table 9.IV – SAWE Inertia Validation ..............................................................................................................................44Table 10.I - Historical Aircraft Tail Volume Coefficients..................................................................................................47Table 10.II - Pitching Moment Coupling with....................................................................................................................51Table 10.III - Rudder Control Power Results for OEI Condition.......................................................................................52Table 10.IV - Longitudinal and Lateral Dynamic Mode Conformity with MIL-8785C ....................................................55Table 10.V – Empennage Surfaces.....................................................................................................................................55Table 11.I - Design Mission ...............................................................................................................................................64Table 11.II - Mission Results..............................................................................................................................................64Table 11.III - Takeoff Results ............................................................................................................................................66Table 11.IV - Landing Results............................................................................................................................................66Table 12.I – Alternate Mission Results ..............................................................................................................................69Table 13.I - Military Vision Specifications ........................................................................................................................72Table 14.I – APU Selection Table......................................................................................................................................74Table 14.II - Fuel System Sizing Requirements .................................................................................................................76Table 14.III - List of Government Furnished Equipment ...................................................................................................77Table 17.I - RFP Compliance Checklist .............................................................................................................................83vii

Nomenclature6DOF Six Degrees of FreedomA 0i Freestream Capture Area, ft 2A 1 Inlet Capture Area, ft 2AB AfterburnerA e Area of Inlet Exit, ft 2AIAA American Institute of Aeronautics andAstronauticsAOA, α Angle-of-Attack, degAPU Auxiliary Power UnitAR Aspect RatioA s Area at Shock, ft 2b W Wing Span, ftCAS Control Augmentation SystemC D Drag CoefficientC f Friction CoefficientCG Center-of-GravityC L Wing Lift CoefficientC Lα Lift Curve SlopeC lδr Roll Moment due to Rudder DeflectionC m Pitch Moment CoefficientC mα Pitch Moment due to AOAC mδr Rudder Pitch Moment CouplingC nδr Yaw Moment due to Rudder DeflectionC yδr Side Force due to Rudder DeflectionDATCOM Air Force Data CompendiumDC Direct CurrentDFSC Digital Flight Control Systemd T Ratio of local pressure to sea-level pressureeSpan Efficiency FactorF F Form FactorGFE Government Furnished EquipmentGPS/INS Global Positioning System/InertialNavigation Systemh AltitudeHUD Heads-up-DisplayICNIA Integrated Communication, Navigation, andIdentification AvionicsIHPTET Integrated High Performance TurbineEngine TechnologyIR InfraredIRSTS Infrared Search and Track System withLaser RangingI x Moment of Inertia about the x-axis, slug-ft 2I xy Moment of Inertia in the xy plane, slug-ft 2I xz Moment of Inertia in the xz plane, slug-ft 2I y Moment of Inertia about the y-axis, slug-ft 2I yz Moment of Inertia in the yz plane, slug-ft 2I z Moment of Inertia about the z-axis, slug-ft 2k 1 Induced Drag FactorK SM Static MarginL/D Lift-to-Drag ratioviiiL/HLANTIRNLBR-TFLEL HTL VTMMMAC, c Wm eMFDMPRLNNACANATONPFNPF cOBIGGSOBOGSOEIPP sP SLqQ FQ TrRAMRATRCSReRFPRTD&ER xR yR zInlet Length to Height RatioLow-Altitude Navigation and TargetingInfraRed for NightLow-Bypass-Ratio Turbo FanLeading EdgeHorizontal Tail Arm, ftVertical Tail Arm, ftMach NumberActual Mass Flow RateMean Aerodynamic Chord, ftEstimated Mass Flow Rate, slug/sMultifunction DisplayMultipurpose Rotary LauncherLoad FactorNational Advisory Committee forAeronauticsNorth Atlantic Treaty OrganizationNet Propulsive ForceCorrected Net Propulsive ForceOn Board Inert Gas Generation SystemOn Board Oxygen Generation SystemOne Engine InoperablePressure, psfSpecific Excess Power, ft/sSea-Level Pressure, psfDynamic Pressure, psfInterference FactorRatio of local temperature to sea-levelstatic temperatureTurn Radius, ftRadar Absorbent MaterialsRam Air TurbineRadar Cross SectionReynolds NumberRequest for ProposalResearch, Test, Development, andEngineeringNon-Dimensional Radius of Gyrationabout the x-axisNon-Dimensional Radius of Gyrationabout the y-axisNon-Dimensional Radius of Gyrationabout the z-axisSurface to Air MissileSociety of Allied Weight EngineersSelf-Contained Energy Storage SystemThrust Specific Fuel Consumption,lbm/lbf-hrSAMSAWESESSFC,TSCFS HT Horizontal Tail Planform Area, ft 2S ref, S W Wing Reference Area, ft 2S VT Vertical Tail Planform Area, ft 2T Thrust, lbf

T Temperature, °RTE Trailing EdgeTJ TurbojetTOGW, Takeoff Gross Weight, lbWT SL Sea-Level Temperature, °RV Velocity, ft/sV 50 Velocity over a 50 ft Obstacle, ft/sVAATE Versatile Affordable Advanced TurbofanEngineV H Horizontal Tail Volume CoefficientVMS Vehicle Management SystemV stall Stall Speed, ft/sV TD Velocity at TouchdownTakeoff VelocityV TOV VW eW fW FW FcVertical Tail Volume CoefficientEmpty Weight, lbFuel Weight, lbFuel FlowCorrected Fuel FlowX x-axisY y-axisZ z-axisβYaw Angle, degη Actual Inlet Efficiencyη rSpec Mil Spec Inlet Efficencyµ brake Braking Coefficient of Frictionµ roll Rolling Coefficient of Frictionρ Density, slug/ft 3ix

1 IntroductionThe American Institute of Aeronautics and Astronautics (AIAA) sponsors annual collegiate design competitions.The request for proposal (RFP) for the 2001-2002 Undergraduate Team Aircraft Design Competition outlinesspecifications for a stealth supersonic interdictor for the US Air Force. The aircraft should be capable of flying a designmission as shown in Figure 1.1. The payload specified for this design mission is shown in Table 1.I. Because multipleweapon loadouts are specified, it is clear that this aircraft, as with any modern aircraft, must have multi-role capabilities.The RFP lists additional requirements for the aircraft including operating constraints and performancerequirements which are summarized in Table 1.II. External tanks may be used but must be retained for the duration ofthe flight. External weapons pylons may also be used suggesting the possibility of a non-stealth configuration. Anotherimportant factor is that the aircraft must have a flyaway cost less than 150 million dollars; a challenging price tag for anaircraft of this size and complexity.Figure 1.1 - Design Mission ProfileTable 1.I - Required Weapons LoadoutLoading # (Quantity) Weapon1 - Design (4) 2000 lb JDAM + (2) AIM-1202 (4) Mk-84 LDGP + (2) AIM-1203 (4) GBU-27 + (2) AIM-1204 (4) AGM-154 JSOW + (2) AIM-1205 (16) 250 lb Small Smart Bomb1

and fly out of the threat area quickly. After retiring the aircraft in 1996, the Air Force decided a new aircraft was neededto drop precision weapons from remote airfields with minimal support.F-15E - “Strike Eagle”The F-15E Strike Eagle (Table 1.III) partially filled the role of the F-111A after it was retired. Although the F-15E airframe was designed for fighter type payloads, it is capable of both air superiority and ground attack missions.Superior maneuverability was achieved with the F-15E due to its high thrust-to-weight ratio and low wing loading.F-117 – “Nighthawk”The F-117 Nighthawk (Table 1.III) also aided in the replacement of the F-111A. However, it has a much lowerpayload capacity and a limited range. The F-117 is also not capable of supersonic speeds and is thus more vulnerable ifit were detected. If a supersonic aircraft were detected, the window of opportunity for an attack is relatively small.Thus, faster aircraft have a tendency to be less vulnerable. Due to the small payload and high maintenance of the firstgeneration stealth technology, the F-117 is a poor replacement for the F-111A.B-1B – “Lancer”The B-1A was designed as a replacement for the B-52. It could carry large nuclear payloads supersonically withan intercontinental range. The SALT treaty limited the B-1 to subsonic speeds and led to the creation of the B-1B. Themission of the B-1A was not unlike that of the RFP proposed mission; however it was performed with the aide of anafterburner in supersonic flight. More information is provided in Table 1.III.B-2 – “Spirit”The B-2 (Table 1.III) is very new to the U.S. inventory. It has intercontinental range unrefuled and carries largeconventional and nuclear payloads. The B-2 is a large aircraft that is very costly to operate.3

Table 1.III - Comparison of the F-111A, F-117, B-2, B-1B, and F-15EManufacturer LockheedGeneralDynamicsBoeing Northrop RockwellDesignation F-117 FB-111A F-15E B-2 B-1BSpan - ft 43.6 32.0 42.8 172.0 78.2Aspect Ratio – – 3 – –Length – ft 66.6 73.5 63.7 69.0 147.0Height – ft 12.5 17.1 18.5 17.0 34.0Wing Area - ft 2 913 – 608 5274 1950Empty Weight – lb 29,500 46,171 32,000 153,700 192,001Payload Weight – lb 5,000 31,500 24,500 40,001 133,999Fuel Weight – lb – – 13,122 200,003 194,999Gross Takeoff Weight – lb 52,501 91,492 81,000 375,998 477,003Max power loading – – 1.73 4.86 –Max Mach # 0.9 2.2 2.5 0 1.25Max combat radius 570 2,750 686 6,300 6,479Service Ceiling - ft – 50,853 – 50,000 –The solution to the RFP is not a trivial design problem. The aircraft will have to be well area-ruled in order tominimize wave drag and have a low frontal radar cross-section. The goal of this design is to meet or exceed RFPrequirements while minimizing manufacturing and operating costs.4

2 Defining the Design DomainAn initial takeoff weight estimate was made using historical aircraft data. First, a database of aircraft similar inmission was compiled. Next, an iterative weight fraction method outlined in Roskam. This method calculates the weightfraction for each mission segment. Using the resulting weight fractions, the aircraft gross takeoff and empty weightswere iterated until a weight fraction consistent with the historical trends was reached. Figure 2.1 shows the historicaltrend in aircraft weight fractions and the initial estimate of Vendetta’s empty and takeoff gross weights. The results ofthe weight fraction method are shown in Table 2.I.Figure 2.1 - Historical Weight Fractions & Weight Fraction EstimatesTable 2.I - Weight Fractions & WeightsMission Segment WeightsWeightsFractionsStart/Takeoff 6% Takeoff 108,400 lbClimb To Cruise 11% Empty 51,600 lbCruise-Out 25% Fuel 47,600 lbDash-Out 17% Payload 9,054 lbDash-Back 14% Fuel Weight Fraction 47.6%Cruise-Back 16%30 Minute Reserve 6%Misc. 5%Total 100%Warm -upMisc.Reserve Initial ClimbCruise BackCruise OutDash BackDash Out5

The weight fraction method provides a rough starting point for aircraft takeoff gross weight. The assumptionsused in the weight fraction method are listed in Table 2.II. Inaccuracies of up to 10% are possible depending on thequality of the initial assumptions, and 20% is not uncommon for unusual missions such as the one outlined in the RFP.Another source of inaccuracy is the historical aircraft used to define weight fraction trends. Because no supercruisingstealth bombers currently exist, many subsonic aircraft or non-stealthy aircraft were used in the historical aircraftdatabase.Once a starting TOGW is known, the physical dimensions can be estimated using a constraint plot. Theconstraint plot examines the relationship between two variables based on given requirements. Generally for initial sizingof an aircraft, the two variables used are sea-level takeoff wing loading and thrust-to-weight ratio. The RFP providesmany performance requirements that can be written as functions of these design parameters. Equations for range,specific excess power, takeoff and landing distance, and others from in Roskam, Nicolai, and Raymer were used todefine the constraints produced by these requirements. Additional assumptions were made to create the constraint plot asshown in Table 2.III.Table 2.II - Weight Fraction AssumptionsTable 2.III - Constraint AssumptionsSFC _CruiseSFC _DashSFC _TurnSFC _LoiterL/D Cruise 10L/D Dash 10L/D Turn 10L/D Loiter 121.11 lbm/lbf-hr1.11 lbm/lbf-hr1.11 lbm/lbf-hr0.8 lbm/lbf-hrC Lmax_TO 1.8C Lmax_CR 1.2C LCruise 0.2AR 3e 0.8The constraint plot in Figure 2.2 shows how thrust-to-weight ratio and wing loading relate to a specificperformance constraint. This allows an acceptable thrust-to-weight ratio and wing loading to be found. Note that anydesign points on the hatched side of a constraint would not meet the specific design requirement that that constraintrepresents. The constraint plot clearly identifies a design domain in which any combination of thrust-to-weight ratio andwing loading would satisfy all of the design requirements.The combination of the weight fraction method and the constraint plot provided an initial estimate the physicalsize and weight of the airplane. From the acceptable wing loading and thrust-to-weight ratio values determined from theconstraint plot, a single point must be chosen. A design with a higher wing loading will result in a smaller aircraft,which will be less expensive and easier to maintain. High thrust to weight ratios will require larger, more expensive andless efficient engines. Many assumptions were used to create the individual constraints, so the true effect of the6

performance requirements on the aircraft is not well defined. Because of this, the initial design point was also chosen inthe center of the design domain in order to allow for aircraft growth. However, more accurate analytical techniques haveallowed the current design point to move closer to the minimum wing loadings and thrust-to-weight ratios. The initialsize and weight estimates are shown in Table 2.IV along with the current values for comparision.1.0Thrust to Weight Ratio0.90.80.70.60.50.40.30.2Current DesignPointRange > 3500 nm2-g P s > 0 fpsInitial DesignPointLanding< 8000 ftInstant Turn Rate > 8°/sec1-g P s > 200 fps1-g P s > 0 fps0.1Takeoff < 8000 ft406080100 120 140Wing Loading, (psf)160180200Figure 2.2 - Constraint PlotTable 2.IV - Initial and Current SizingInitial Design Point Current Design PointTOGW 108,400 lb 125,051 lbW e 51,600 lb 56,797 lbW f 47,600 lb 58,974 lbW f /W 0.43 0.47T/W 0.54 0.48W/S 100 lb/ft 2 83.3 lb/ft 2T SL 58,000 60,000 lbS ref 1,084 ft 2 1,500 ft 27

3 ConfigurationThe current configuration of the aircraft is the result of several major iterations. The first iterations wereindividual designs developed by each of the six team members at the beginning of the project. Four of the individualdesigns were considered for the configuration the Vendetta.The Nergal (Figure 3.1) and the Jackhammer (Figure 3.2) were both tailless aircraft utilizing thrust vectoring forstability in the yaw axis. Each aircraft utilized Pratt and Whitney F119 engines found in the F-22. The Nergal was farsmaller than the Jackhammer due to its single engine and fuel volume usage. Both aircraft incorporated rotary weaponslaunchers capable of carrying every weapon mentioned by the RFP in a single weapons bay. The cockpit layout wasside-by-side to minimize the redundant cockpit displays and maximize crew communication.The Interdictor (Figure 3.3) and Big Paulie (Figure 3.4) were both based on the RFP engines. The Interdictorused straight inlet ducts with an inlet screen similar to what is used on the F-117. Big Paulie attempted to make use ofthe axisymetric translating center body inlets of the RFP engine in a stealth design by using a prismatic translating inletspike to better control radar energy.Both aircraft were deemed impractical due to the large cross-sectional areasproduced by the excessive size of the RFP engines. The cockpit arrangement of these aircraft was side-by-side similar tothe Nergal and Jackhammer.Analysis performed during this early work clearly showed that the engine provided by the RFP was far too largefor the thrust it provided. The two aircraft designed for the F119 were both smaller and more space efficient. Thisnarrowed the design options to the Nergal and Jackhammer. Both of these aircraft were tailless and it was determinedthat the weight and drag benefits associated with the lack of a vertical tail would be outweighed by the costs associatedwith the thrust vectoring system. It was also determined that the aircraft were too unstable laterally to be controlled byan inexpensive, low bandwidth, thrust vectoring system. It was decided to begin with a new design incorporating thestrong points of each aircraft.The first iteration of the aircraft is shown in Figure 3.5; it is a large aircraft that has many design flaws. The firstFigure 3.1 - Nergal Figure 3.2 - Jackhammer Figure 3.3 - Interdictor Figure 3.4 - Big Paulie8

and most obvious is the above-chine mounted inlet, easily seen in the front view. The chine causes a vortex roll-up thatwould be directly ingested by the inlet at moderately high angles-of-attack (AOA). A low bypass ratio engine mighttolerate these flow disturbances without problems; however, the design utilizes a new engine with a bypass ratio ofapproximately 1.5. This type of engine will not tolerate swallowed vortices.Another problem with the initial configuration was weight distribution. The fuel center-of-mass was not near theempty weight center-of-mass. This caused the aircraft to take off very stable and land very unstable. This could not beremedied due to the small volume available for fuel in the aft portion of the fuselage. The majority of the fuel volume inthe aft portion of the aircraft was located around the engines. This is undesirable due to the possibility of a catastrophicfailure of the engine fan disk or afterburner.Another problem arises from the 20° facet on the bottom of the fuselage. This created a large radar footprintunderneath the aircraft, as shown on the right side of Figure 3.5. The vertical stabilizer also created poor low observablecharacteristics. The final flaw that drives the aircraft to the new configuration is the pitching moment characteristics ofthe fuselage. The side-by-side seating arrangement of the first iteration caused the fuselage to be excessively large in theareas forward of the aircraft’s center-of-gravity. The pitch up tendencies of the aircraft grew very large with only smallAOA. The control power of the horizontal surfaces was found to be incapable of combating the problem.The second configuration shown in Figure 3.6, shows significant design evolution from the previousconfiguration. This configuration features many changes that aid in solving the previously discussed problems. Thecockpit was changed to a tandem arrangement the single vertical tail was replaced by twin canted surfaces. The engineswere moved to the top of the fuselage to avoid detection from infrared sensors. The takeoff gross weight decreased to114,000 lb due to improved engine and aerodynamics estimates.• Span = 50 ft• m.a.c. = 23 ft• S ref = 965 sq. ft• TOGW = 121,600 lb• Empty Weight = 62,000 lb204060dBm 250’19’105’23’Figure 3.5 - Initial Configuration9RFPRequirement

• Span = 53 ft• m.a.c. = 32 ft• S ref = 1500 sq. ft• TOGW = 114,000 lb• Empty Weight = 55,000 lb53’19’98’35°40 dBm 23020RFPRequirement18’13°Figure 3.6 - Second ConfigurationThe design approach for the second configuration was differed significantly from that the first. The center-ofgravitywas decided on before the first part was placed on the aircraft and every effort was utilized to keep it in theappropriate place. The weight and balance issues, though still present, were dramatically improved. The fuel load andpayload compartment reside directly on the desired center-of-gravity; however, the empty weight center-of-gravity wastoo far aft. The low mounted wing proved to be a structural challenge when incorporating a landing gear well. Anotherissue dealt with the cruise AOA. It was shown that the aircraft would cruise at approximately 4 degrees. The forwardchine on the fuselage would be shedding a vortex throughout the cruise portion of the mission resulting in higher drag.The chine angle should meet the onset flow angle. The maximum radar signature of the aircraft decreased dramatically(by 10dB) from the previous configuration however, the radar return in the frontal aspect increased substantially from-12dB to 0dB. The frontal aspect is an important design requirement thus another revision to the aircraft was createdpaying more attention to frontal RCS.The final configuration of Vendetta is shown in Foldout 1. The configuration was generated with the samemethodology as the second iteration; however, greater attention was given to load paths and landing gear placement.The 4° cruise AOA was incorporated into the forward chine. The Vendetta has grown a small amount and currentlyweighs 125,000 lb. The aircraft has a tandem cockpit supported by a very long nose. The long nose offsets the mass ofthe large engines and the massive structure required for the full flying horizontal stabilizers. The engines can beremoved through the bottom of the aircraft, as there are no primary load paths obstructing access. This makesmaintenance easier for the ground crew. The APU is located in the engine compartment keeping the fuel and fireretardant systems as redundant as possible. The inlets are under-wing mounted to keep them in clean flow throughoutthe flight envelope. The Vendetta has a 1500 ft 2 wing area with a leading edge sweep of 40 degrees. The design driverswill be discussed in detail throughout following sections. The inboard layout can be seen in Foldout 2.10

Weight BuildupComponent Weight Fuselage Butt PlaneStation (ft)(ft)StructuresWaterLine (ft)Wx(lb-ft)Wy(lb-ft)Wz(lb-ft)Wing Group 8,779 66.0 0.0 -3.0 579,319 0 -26,511Horizontal Tail 1,262 94.3 0.0 -3.0 119,073 0 -3,812Vertical Tail 1,279 86.2 0.0 -7.1 110,176 0 -9,117Fuselage 10,540 44.2 0.0 -2.1 465,770 0 -22,345Main Gear 2,289 68.9 0.0 -1.9 157,648 0 -4,371Nose Gear 400 17.3 0.0 4.7 6,915 0 1,87524,548 58.6 0.0 -2.6 1,438,902 0 -64,282PropulsionEngines 11,034 84.9 0.0 -2.7 937,198 0 -29,791Engine Mounts 138 84.9 0.0 -2.7 11,752 0 -374Firewall 102 84.9 0.0 -2.7 8,638 0 -275Nozzle 140 84.9 0.0 -2.7 11,892 0 -378Oil Cooling 77 84.9 0.0 -2.7 6,528 0 -208Starter 185 84.9 0.0 -2.7 15,679 0 -498SystemsPressurization11,67537284.917.30.00.0-2.7-4.0991,6876,43300-31,523-9,395Air Induction 2,325 64.4 0.0 -3.2 149,738 0 -788Anti-Ice 248 17.3 0.0 -4.0 4,289 0 -1,414APU 350 80.9 0.0 -4.3 28,326 0 -2,509Auxillary Gear 578 17.3 0.0 -4.0 10,005 0 -4,531Avionics 1,122 17.3 0.0 -4.0 19,416 0 -1,923C.G. Control System 476 44.2 0.0 -2.1 21,037 0 -2,917Electrical 1,376 44.2 0.0 -2.1 60,811 0 -137Engine Controls 65 84.9 0.0 -2.7 5,507 0 -10,130Flight Controls 3,752 17.3 0.0 -4.0 64,945 0 -2,492Fuel System 5,101 58.3 0.0 -2.1 297,236 0 -10,934Furnishings/Equipment 617 17.3 0.0 -4.0 10,675 0 -10,100Launchers & Weapons 2,500 55.3 0.0 -1.8 138,300 0 -1,965Hydraulic System 1,110 44.2 0.0 -2.1 49,052 0 -60Oxygen System 28 17.3 0.0 -4.0 486 0 -2,242Paint 555 44.2 0.0 -2.1 24,526 0 -43,61820,574 43.3 0.0 -5.1 890,782 0 -105,155Geometric DataItemAreasUnits Wing Horizontal VerticalReference sq. ft 1500 270 165Exposed sq. ft 900 265 160Wetted sq. ft 1714 528 330Span ft 54.8 35.1 9.2Aspect Ratio - 2.0 4.6 2.1Taper Ratio - 0.17 0.23 0.17SweepsLE ° 40 40 40c/4 ° 21 24 21c/2 ° 5 9 5TE ° -30 -30 -30ChordsMean Aerodynamic ft 32.0 14.9 10.7Root ft 46.8 15.0 16.0Tip ft 8.0 3.5 2.8BP 0.0Total Fuselage Volume: 5000.8 cubic feetFS295.6Forward FuselageBreak Point 368.4" MAC @ BP 125Wing Break PointProjected Vertical PlanformAft FuselageBreak Point178.0" MAC@ BP 11153°EMPTY 56,797 58.5 0.0 -3.5 3,321,371-200,959PayloadGBU-27 2,000 55.3 0.0 -1.8 110,640 0 -3,540GBU-272,000 55.3 0.0 -1.8 110,640 0 -3,540GBU-272,000 55.3 0.0 -1.8 110,640 0 -3,540GBU-272,000 55.3 0.0 -1.8 110,640 0 -3,540AM-120 390 55.3 0.0 -1.8 21,575 0 -690AM-120 390 55.3 0.0 -1.8 21,575 0 -690Crew & Cargo 500 55.3 0.0 -1.8 27,660 0 -8859,280 55.3 0.0 -1.8 513,370 0 -16,426054.8' (657.3")35.1' (421.8")13.5' (161.6" )20° CantFS122.7FS226.2FS275.6FS303.4FS362.8FS413.3FS551.1FS631.5FS688.9103.3' (1240.0")FS729.414°FS826.7FS964.5FS940.0128.3" MAC @ WL 101FS1051.3FS1102.3FS1175.8ZERO FUEL 66,077 58.0 0.0 -3.3 3,834,740 0 -217,38520.6'(247.5")FuelFwd. Fuselage 23,034 41.6 0.0 0 -1.9 958,460 0 -43,996Left Wing 6,366 67.0 12.9 -3.0 426,331 82,249 -19,162Right Wing 6,366 67.0 -12.9 -3.0 426,331 -82,249 -19,162Aft FuselageTAKEOFF GROSS23,20858,974125,05170.058.358.10.00.00.0-1.9-2.1-2.71,625,2333,436,3557,271,096000-44,095-126,414-323,79915.6' (187.0")F.R.L @ WL 0135°fov5.3'(63.4")VendettaChris DroneyNate Schnaible13°Kolby KeiserChris MaglioFoldout 1Aircraft OverviewChris AtkinsonDan SalluceRev. 3Scale 1:150High Rollers5/23/02

VendettaChris DroneyNate SchnaibleKolby KeiserChris MaglioFoldout 2Inboards and SectionsChris AtkinsonDan SalluceRev. 3Scale 1:150High Rollers5/23/02Retracted Main GearFull Flying StabRadar and NoseAccessoriesEngineRetracted Nose GearBP 30.0Partial Section - BP 30.0BP 0.0Forward Fuselage Tank23,034 lb JP-8Total Fuel58,974 lb JP-8Aft Fuselage Tank23,208 lb JP-8Left Wing Tank6,366 lb JP-8Tandem CockpitWeapons BayInletAPUFS295.628°30°FS122.75°FS226.2FS275.6FS303.4FS362.8FS413.3FS551.1FS631.5FS688.9FS729.4FS826.7FS964.5FS940.0FS1102.3FS1051.3FS1175.8Radar and NoseAccessoriesRetracted Main GearEngine AccessoriesFS122.7FS226.2FS275.6FS295.6FS303.4FS362.8F.R.L11°AN/APG-77IRSTAAR PortWeapons BayFlow DeflectorMPRLAPUWing Tank X 26,366 lb each70% Volume UsageNo Fuel Around Engineor Near Fan FaceForward Fuselage Tank23,034 lb JP-8BP 0.0Right Wing Tank6,366 lb JP-8Left Wing Tank6,366 lb JP-8FS413.3FS551.1FS631.5FS688.910° Fall AngleFS729.4FS826.7Aft Fuselage Tank23,208 lb JP-8Fuel Tank ContainedWithin Section BreaksForward Fuselage Tank23,034 lb80% Volume UsageAft Fuselage Tank23,208 lb80% Volume UsageTotal Fuel58,974 lb JP-8FS940.0FS964.5FS1051.3FS1102.3FS1175.8

4 Stealth ConsiderationsRadar cross-section (RCS) is an important low observability consideration for the Vendetta. The geometricshaping of an aircraft is the main contributor to its radar return. When radar energy interacts with the surface of anaircraft many phenomenon affect the resulting disturbance to the electromagnetic energy. Radio waves that strike asurface may reflect off of that surface or begin to travel along the surface. When edges are encountered, energy is eitherradiated outward in planes perpendicular to the edge or reflected back along the surface. To achieve low radar crosssectionin any particular aspect of an aircraft, the surfaces of the aircraft must be shaped so that the electromagneticenergy is either absorbed, or reflected away from the receiving station. After shaping, radar absorbing materials (RAM)can be utilized to minimize the spikes created by problem areas such as inlets, wing tips, and control surfaces.The design features described below and illustrated in Figure 4.1 are used to control the radar returns in specificaspects. The fuselage is constructed from flat sides and constant radius curves to produce radar returns in a singledirection away from the source of the radar energy. The sides are kept at a 60° angle from the horizontal and the bottomis kept flat to minimize the radar footprint that is created below the aircraft. The vertical tails are canted to avoidcreating perpendicular surfaces which would return radar energy directly back to its source. The leading edge sweep is40° creating spikes well off of the frontal aspect of the aircraft. All other leading edges are kept swept at this same angleto concentrate radar energy into the same regions. An analysis of radar threats, as shown in Foldout 3, indicates thatmost Vendetta will require low signatures from frontal aspect required by the RFP to a 15° look up angle.The RFP specifies that the aircraft incorporate balanced observables. Infrared (IR) sensors present anotherobservability threat. Emissivity matching will be employed to minimize the infrared energy radiated from hot surfaceson the aircraft. Specially designed paints and surface treatments will be used to match the emissivity of the aircraft tosurroundings, aiding in the disappearance of the aircraft to any IR sensors. As will be shown in the propulsion section, a40° LE SweepAll other SurfacesMatchedHidden CantedVerticalsFigure 4.1 - Stealth Considerations60° Facet13

low signature axisymmetric advanced nozzle will be used that has been developed for use on low observable aircraft.Visual observability will be addressed through the use of mission planning and contrail avoidance. No practical visualstealth technology currently exists that could be incorporated into the design aside from color selection.To quantitatively analyze the radar cross-section of the Vendetta, Radbase2 software by Surface Optics wasutilized. First, a faceted model was generated from the 3D model. Faceting was limited to only those necessary becauseof the demanding processing requirements. Facets were limited to 10 degree tolerances at roughly 0.017 feet minimums.The facetted model is presented in Foldout 3. It can be seen that heavy facet optimization was needed to make sure thatall facets met tangency requirements to leave smoothly curved and splined surfaces. The spline arc on the top of thefuselage is modeled with facets every 10°. For the flat surfaces like the wings and empennage 10° angular spacing ismore than adequate to represent the surface.The Radbase2 RCS code calculates radar returns based on Physical Optics and Chu-Stratton integral methodswhich are computationally intensive. Because of this, bounces off surfaces were limited to two after the initial bounce.Vertical-vertical return and transmission polarization were analyzed as it is the most relevant to how radar stationsoperate. Horizontal-horizontal as well as mixed HV and VH returns did not yield significant returns. Monostatic radarswhich both broadcast and receive radar waves were used in the analysis. Although Radbase2 can calculate bistaticreturns, there are literally an infinite number of threat situations possible and the RFP does not specifically call out arequirement.The code was allowed to iterate on the model with 1° azimuth increments and for 0° and 15° lookup angles. Itwas also run for 1, 5, 10, and 12 GHz radar frequencies. Most fast track and search radar runs at the higher frequencieswhile long range threat radars utilize the lower frequencies. The 1 to 10 GHz range covers most of the radars that areexpected for the role of this aircraft and are specifically required by the RFP. A table of common ground and surfaceradars with their respective frequencies is presented as in Foldout 3.Although the information for common radars is available for those currently used by the United States, radarenergy and the principles of their propagation through air are similar regardless of application. Looking at the radarsused by the Navy shows that lower frequency radars are better suited to traveling longer distances with largerwavelengths. Fast track radars are more suited for higher resolutions and fast, short range surface to air missiles (SAMs).Data are not readily available for radars made by foreign manufacturers.The 1 to 12 GHz range covers FM and XM radar bands which are the most common threats. The RFP specificallyrequires that the Vendetta have a frontal RCS of 0.05 m 2 in the 1-10 GHz range. As the threat chart shown in Foldout 314

shows, most threats will be from below and at shallow angles of about 15° while at 50,000 ft during ingress. Because ofthis, the 0° and 15° lookup angles were analyzed. The results of the Radbase2 software are illustrated first in Foldout 3.which depicts the radar cross-section of the aircraft from a frontal, or 0° lookup angle.The figure shows that the vehicle does clearly meet the frontal RCS requirement of 0.05 m 2 (-13 dBm 2 ) set forthin the RFP. It also shows that the iterative measures taken to shape the aircraft worked. The leading edge and trailingedge of the wing come together closely. There is a large return directly from the side of the aircraft due to the wingtipand fuselage side. It can also be seen that, although there are slight variations in the returns due to the differentfrequencies, they do not vary much due to the fact that the Vendetta is a rather large vehicle; hence none of the surfacesare small enough to interfere at the radar wavelengths. The weakest azimuth angle for the Vendetta is the 40° anglewhere the leading edge sends a large spike forward. However, the Vendetta meets the RFP requirement for a full 77° ofazimuth.Looking at the equally crucial 15° lookup angle cross-section in Foldout 3 reveals a slightly different picture. Itshows that the Vendetta meets and exceeds the 0° lookup angle returns. This is highly advantageous. The shape of thebottom of the aircraft is effective in keeping spikes at a minimum. As mentioned earlier, this is a crucial area for theVendetta. As most of its threats are from the ground, it is important that the aircraft has a limited return in thisorientation.The threat chart shows that the Vendetta would remain in range of the Soviet SA-12 and SA-6 SAMs for 160seconds and 60 seconds, respectively. This means that the returns from the bottom of the Vendetta are crucial for threatassessment. The software was used to generate an RCS butterfly plot in a sweep around the vehicle to determine thefootprint that it will leave as it flies above its threats. Foldout 3 shows this sweep.It can be seen that the 60° facets on the bottom of the fuselage are deflecting radar away from the vulnerablelookup orientation. The aircraft is still producing a large return of almost +40 dBm 2 in this position, however. Onceagain, there is little variation in the returns for various frequencies. Mission planning would become crucial to be surethe Vendetta avoids flying directly over only these long-range, high flying threats such as the SA-12.It is important to note that the addition of radar absorbing material (RAM) would further reduce some of thereturns on the aircraft. Note that all plots shown reflect the fact that software is assuming fully reflective metal on allsurfaces. No cavities are being modeled besides the inlets. This is a conservative approach. RAM could be applied inactuality to reduce some of the returns on the bottom and front of the aircraft.15

0° Lookup77°50403020100-10-20-30-40Threat Frequency Analysis:RadarAN/TPS -43E Mobile RadarAN/TPS -70Fixed Ground RadarManufacturerWestinghouseNorthrop GrummanFrequency2.9 to 3.1GHz2.9 to 3.1GHz0AN/SPS -49850 to 942Typical Long Range Naval Navy Research LabsMHzRadar1 GHz5 GHz10 GHz12 GHzRFP RequirementRegion of RFPAN/SPS -559.05 toCompliance -13 dBm 2 Long Range Surface Search ISC Cardion10.0 GHzRadarImageRadial Sweep6050403020101 GHz5 GHz10 GHz12 GHz15° Lookup3063°Threat Envelope Analysis:100,000 ftSA-1220100-10-20-30-4075,000 ft25,000 ft1,000 ft50,000 ft8°15°SA-6SA-11SA-8/15SA-9/13500 ft1 GHz5 GHz10 GHz12 GHzRFP RequirementRegion of RFPCompliance -13 dBm 2300 ft200 ft45706040503020 nm40 30 20sm10Time In RangeSA-12 -- 160 sSA-6 -- 60 sSA-11 -- 55 sSA-9SA-13SA-8SA-11SA-13SA-6SA-12Chris DroneyNate SchnaibleRev. 3VendettaKolby KeiserChris MaglioHigh RollersFoldout 3Low ObservablesChris AtkinsonDan Salluce5/23/02

5 Aerodynamics5.1 Wing PlanformThe first aerodynamic parameters that were considered were the wing planform area and aspect ratio. To selectthe optimum wing planform area and aspect ratio, the effect of these two parameters on the specific excess power (P s )and fuel consumption over the design mission were studied. The 1-g military specific excess power at an altitude of50,000 ft and Mach number of 1.6 was estimated using engine data coupled with drag estimation based on componentskin friction drag and area ruling. Fuel consumption was estimated by numerically integrating engine fuel flow over thedesign mission. The additional weight and maximum cross-sectional area of larger wing areas were considered incalculations; however, the mission profile and fuel weight at takeoff were kept constant. The results shown in Figure 5.1indicate that a wing planform area of approximately 1,500 ft 2 and aspect ratio of 2.0 would maximize specific excesspower and minimize fuel consumption.Specific Excess Power (ft/s)Mach 1.6, 50,000 ft, and Maneuver Weight969492908886841,600 ft21,800 ft 22,000 ft 2Design Point1,400 ft 21,200 ft 2Wing Area2.52.252.01.75 1,000 ft 22,200 ft 2 Aspect Ratio1.52,400 ft 28257 58 59 60 61 62 63 64 65 66 67Fuel Burn over Mission (1,000 lb)Figure 5.1 - Optimization of Wing Area and Aspect Ratio17

5.2 Wing SweepThe next wing parameters considered were the leading and trailing edge sweep angles. Because any edges on anaircraft reflect radar energy, the sweep angles of the wing were chosen to minimize radar energy reflected back to thesource, especially in the frontal aspect of the aircraft where a specific RCS requirement is given by the RFP. To avoidreflecting radar toward the front of the aircraft, the leading and trailing edges of the wing had to be highly swept. Inaddition, 45º sweep angles could not be used because a corner reflector would be created reflecting radar energy back toits source from any direction. These requirements led to a diamond shaped wing planform with a leading edge wingsweep of approximately 40º. Two initial designs were considered one having a 40º swept leading edge and a 30º forwardswept trailing edge and the other having matched 35.3º leading edge and trailing edge sweeps. A trade study wasperformed to select between these two wing configurations by studying the effect of the two configurations on RCS andaerodynamics. Figure 5.2 shows a comparison of radial sweeps of the aircraft with both configurations using RadBase2.The return from the 40º and 35.3º leading edge sweeps can be clearly seen in the plot. The leading edge spike on thematched leading and trailing edge configuration is approximately 15 dBm 2 lower than the other configuration, however,it is 5º closer to the frontal aspect of the aircraft. The aerodynamic study of the two wing configurations indicated thatapproximately 1,500 lb of additional fuel would be required due to the additional wave drag from the lower leading edgesweep angle. Because of the aerodynamic benefits of a higher leading edge sweep angle, and because the RFP onlygives frontal aspect RCS requirements, the 40º leading edge and 30º trailing edge configuration was chosen.Once the optimum wing area, aspect ratio, and sweep angles were identified, the tip chord was kept at 8 ft toavoid overly small tip chords that could interact unpredictably with radar wavelengths. This resulted in the wingplanform shown in Figure 5.3, with the measurements given in Table 5.I. Leading edge flaps, trailing edge flaps, andailerons were added to the wing. The chords of the high lift devices and control surfaces were kept at a constantpercentage of the mean aerodynamic chord so that hinge lines would parallel the wing edges and would not createadditional RCS spikes. The trailing edge flap chord is 20% of the mean aerodynamic chord and the aileron and leadingedge flap are each 10% of the mean aerodynamic chord. The trailing edge flap extends from the fuselage to 65% of thesemi-span, the leading edge flap extends from the fuselage to 90% of the semi-span, and the aileron extends from theedge of the flap to 90% of the semi-span. No moveable surfaces were added to the last 10% of the semi-span so thatradar absorbing materials could be added in the wing tip to minimize any returns from that edge.18

50 dB40 dB30 dB20 dB10 dB0 dB-10 dB-20 dB-30 dB-40 dB-50 dB35.3º LE Sweep40º LE Sweep1 GHz. 40º LE Sweep10 GHz. 40º LE Sweep1 GHz. 35.3º LE Sweep10 GHz. 35.3º LE SweepRFP Requirement (-13 dB)Figure 5.2 - Effect of Wing Leading and Trailing Edge Sweep on RCSFigure 5.3 - Wing PlanformTable 5.I - Wing MeasurementsPlanform Area 1,500 ft 2Span54.8 ftRoot Chord46.8 ftTip Chord8.0 ftMAC32.0 fty Location of MAC 10.5 ftAspect Ratio 2.0Leading Edge Sweep 40.0ºSweep at Quarter Chord 20.5ºSweep at Half Chord 4.7ºTrailing Edge Sweep -30.0ºTaper Ratio 0.17Leading Edge Flap Area 112 ft 2Trailing Edge Flap Area 137 ft 2Aileron Area 44 ft 2Flapped Wing Area 624 ft 219

5.3 Wing ThicknessThe effect of wing thickness-to-chord ratio (t/c) on performance was studied so that the optimum t/c could bechosen. Initially, a wing thickness of 3% of the chord was chosen based on existing supercruise aircraft. Increasing theroot thickness of the wing was considered to reduce the weight of the wing and to increase fuel volume in the wing. Theeffects of wing root thickness on wing weight, cross-sectional area, fuel consumption, and fuel volume were studied.The weight of the wing was estimated using the method presented in Raymer, and the additional cross-sectional area wascalculated numerically. The resulting wing weights and cross-sectional areas for wing root t/c from 3% to 6% are shownin Figure 5.4. The effects of the resulting weights and cross-sectional areas on the fuel consumption during the missionwere estimated using the same method used for the wing sizing. The results in Figure 5.5 show that the additional fuelconsumption over the mission due to the wave drag of a thicker wing greatly exceeds the additional fuel capacity of thewing. A constant wing t/c of 3% was chosen because it minimizes both wave drag and fuel consumption over themission.9,000t root = 3%8,500Weight of Wing (lb)8,0007,500t root = 4%t root = 5%7,000t root = 6%6,50013 14 15 16 17 18 19 20 21Maximum Frontal Cross Sectional Area of Wing (ft 2 )Figure 5.4 - Effect of Root Thickness-to-Chord Ratio on Wing Weight and Cross-Sectional Area20

80,000Fuel Consumption over Mission (lb)75,00070,00065,00060,00055,000Fuel Onboard50,0003.0% 3.5% 4.0% 4.5% 5.0% 5.5% 6.0%Wing Root ThicknessFigure 5.5 - Effect of Root Thickness-to-Chord Ratio on Fuel Consumption5.4 AirfoilThe NACA 65A-003 airfoil section was chosen because a symmetrical airfoil is optimum for supersonic flight.The airfoil ordinates given in Theory of Wing Sections for an NACA 65A-006 were scaled and interpolated usingLagrangian polynomials to define the geometry of the wing. The leading edge radius of the airfoil is 0.1% of the chord,which is approximately 0.375 inch at the mean aerodynamic chord and 0.100 inch at the tip. The airfoil sections at themean aerodynamic chord and tip of the trailing edge flap are shown in Figure 5.6 and Figure 5.7, respectively. Becausethe chords of the flaps remain constant as the wing chord changes, each airfoil section has a different relative flap sizes.0.10.050-0.05-0.10 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1Figure 5.6 - Airfoil Section at MAC21

0.10.050-0.05-0.1-0.150 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1Figure 5.7 - Airfoil Section at Tip of Trailing Edge Flap5.5 Lift CurveThe lift curve slope of the wing was first estimated using standard subsonic theory, compressibility corrections,and linear supersonic theory. Next, the stall angle-of-attack of the wing was estimated under takeoff and landingconditions by calculating the lift distribution of the wing using LinAir. The section lift coefficient was calculated as afunction of the spanwise location of the section for different wing angles-of-attack. The wing was assumed to stall whenone of the section lift coefficients exceeded the maximum lift coefficient given in Theory of Wing Sections. The wingstall angle-of-attack was determined to approximately 14º. Because the wingtip was shown to stall at a much lowerangle-of-attack than the rest of the wing, adding a –3º angle of incidence to the wingtip was considered. The resultingtwist extends the stall angle-of-attack to approximately 16º; however, the twist decreased the lift coefficient at a givenangle-of-attack and could impact RCS and supersonic aerodynamics. Ultimately, the non-twisted wing was chosen,because with the use of a leading edge flap, the additional angle-of-attack range was not needed. The lift distributions ofthe wing with and without twist are shown in Figure 5.8. The lift curve slope and lift distributions were also investigatedwith PanAir (Boeing Code A502i). LinAir only models the vortex lattice produced by a given planform, whereas PanAirmodels wing shape and thickness using sources and doublets. Because wing thickness was modeled, PanAir predictedthe wing lift curve slope to be slightly higher than LinAir.The effects of the trailing edge flap were estimated using the stall angle-of-attack and lift coefficient incrementsgiven in Nicolai. The effect of the leading edge flap was estimated by assuming that a 10º leading edge flap deflectionwould increase the stall angle-of-attack by approximately 10º, and the decrease in lift coefficient was estimated based onthe change in effective angle-of-attack of the airfoil sections. The resulting subsonic lift curve at Mach 0.2 is shown inFigure 5.9 along with the lift curve slopes predicted by LinAir and PanAir.22

1Section Lift Coefficient0.90.80.70.60.50.40.3Max. Section Lift Coefficient0º Tip IncidenceAOA16º15º16º15º14º14º 13º 13º 12º 12ºAOA-3º Tip IncidenceCalculated Using LinAir0.20.100% 10% 20% 30% 40% 50% 60% 70% 80% 90% 100%Spanwise Distance (percent semi-span)Figure 5.8 - Lift Distribution of Wing with and without Twist2Lift Coefficient1.510.50-0.5C L = 1.16C L = 1.15C L = 0.53C L = 1.51C L α = 2.33 1/rad10º LE FlapDeflection30º TE FlapDeflection PanAirLinAirCleanTail Strike Angle (13º)-1-20 -15 -10 -5 0 5 10 15 20 25 30Angle-of-Attack (degrees)Figure 5.9 - Subsonic Wing Lift Curve (Mach 0.2)23

5.6 DragDrag was divided into four parts: parasite drag, wave drag, induced drag, and trim drag. The parasite dragcoefficient was estimated using a component buildup method with form and interference factors as shown in Table 5.II.The wave drag was calculated using the formula presented in Brandt & Stiles. The wave drag efficiency factor wascalculated from cross-sectional area distributions using the de Kármán integral and the theoretical wave drag of a perfectSears-Haack body. The cross-sectional area distributions were measured at transonic (Mach 1.0) and supersonic (Mach1.6) conditions. The transonic area distribution was measured by passing vertical planes through a solid model of theaircraft and measuring the intersecting area. The supersonic area distribution was measured by passing Mach conesthrough the model, measuring the intersecting area, and projecting that area onto the vertical plane. For both cases, theengine capture area was subtracted from sections containing the inlet, engine, and nozzle. The resulting areadistributions shown in Figure 5.10 and Figure 5.11 both match reasonably well with that of a perfect Sears-Haack body.Both distributions yield a wave drag efficiency factor of approximately 2.14 (based on 80 ft 2 max. area and 100 ftlength).9080WingCross Sectional Area (ft 2 )706050403020Sears-HaackFuselageVertical TailHorizontal Tail1000 200 400 600 800 1,000 1,200Fuselage Station (inches aft datum)Figure 5.10 - Transonic Area Distribution (Mach 1.0)Induced drag was estimated using standard subsonic theory and the supersonic equation presented in Brandt &Stiles to calculate the induced drag term (k 1 ). Trim drag was calculated as induced drag generated by the horizontal tailat the lift coefficient required to trim the aircraft with a given static margin and zero lift moment coefficient. Theresulting drag build-up at an altitude of 50,000 ft, maneuver weight, and 5% static margin is shown in Figure 5.12.24

9080WingCross Sectional Area (ft 2 )706050403020Sears-HaackFuselageVertical TailHorizontal Tail1000 200 400 600 800 1,000Fuselage Station (inches aft datum)Figure 5.11 - Supersonic Area Distribution (Mach 1.6)Table 5.II - Parasite Drag Component Buildup (50,000 ft, Mach 0.5)Component Wetted Area Length Re F F Q F C f C DFuselage 2,500 ft 2 100.0 ft 59,060,935 1.09 1.10 0.00224 0.00446Wing 1,714 ft 2 32.0 ft 18,878,786 1.22 1.00 0.00266 0.00370Horizontal Tail 528 ft 2 15.0 ft 8,859,140 1.22 1.08 0.00299 0.00139Vertical Tail 330 ft 2 10.7 ft 6,319,520 1.22 1.08 0.00316 0.00092Σ 0.010470.060.05Drag Coefficient0.040.030.02300.02Induced DragTrim Drag0.01Wave DragParasite Drag00 0.5 1 1.51.62 2.5 3MachFigure 5.12 - Drag Build-Up at 50,000 ft, Maneuver Weight, and 5% Static Margin25

6 PropulsionIn developing the propulsion system for the Vendetta, the RFP specifications of supersonic cruise and stealth arethe driving factors. Due to the frontal RCS requirement, the fan blades of the engine must remain hidden which drivesthe engine placement well inside the aircraft.6.1 Engine SelectionThe RFP specifies a Low-Bypass-Ratio Turbofan (LBR-TF) or a Turbojet (TJ) engine may be used to perform themission. Both sizing equations and a candidate engine deck, an axisymmetric center body inlet and a mixed flow ejectornozzle, were supplied with the RFP, with an option to use either or neither. Since it included physical dimensions andfuel flow values the RFP engine deck was used instead of the equations provided by the RFP. The RFP enginespecifications are shown in Table 6.I.Table 6.I - Engine Specifications of RFP Supplied EngineEngine and Nozzle LengthPropulsion System LengthFan Face DiameterMaximum DiameterWeight with Nozzle310 in425 in50 in65 in7200 lbThe engine supplied by the RFP includes fuel flow and thrust data for part power, idle power, and military power.All engine data supplied by the RFP are corrected to sea level and a Mach number of zero. Therefore, every value forthrust and fuel flow at each altitude and Mach number is given in corrected net propulsive force (NPF c ) and correctedfuel flow (W Fc ). To find the actual thrust (NPF) and fuel flow (W F ) the following equations were used:NPF = NPF ⋅ dcT0.6F=F⋅c T⋅TW W Q ddT= +MPP2 3.5(1 .2 ) ( )SLT2QT= 1+0.2 M ( ) TSLOnce the data were uncorrected the military thrust was found. The RFP supplied equations that could be used to scale theengine based on a desired thrust. The scaling equations are as follows:26

NPFNewMeasurement = OldMeasurement( ) NPFAxial length scaling exponent = 0.4Diameter scaling exponent = 0.5Weight scaling exponent = 1.0baseexponentThe RFP engine produced a military thrust of 26,350 pounds and had a cruise thrust specific fuel consumption(TSFC) of 1.19 1/hr for Mach 1.6 flow at 50,000 ft. TSFC is calculated using the following equation:TSFC =W FNPFThe Vendetta will require two engines to perform the desired mission. The size, weight, and location of theengines have great effect on the size of the airplane. The larger the engines the wider the aft portion of the fuselage andthe longer the airplane. For the size and weight of the RFP engine, it produced too little thrust and burned too much fuelcompared to modern turbofan engines.Other engines were analyzed in an attempt to find a better performing engine that was smaller and lighter thanthat supplied. Through this research the Concorde Rolls-Royce SNECMA Olympus engine was found to be comparableto the RFP engine; however, the engine was first manufactured and flown in the Concorde in the mid ‘60’s through mid‘70’s. Table 6.II compares the RFP engine to that of the SNECMA Olympus. As can be seen, the SNECMA Olympus isvery close in size and weight to that of the RFP; however, it produces even more thrust than that of the RFP. Also theweight of the SNECMA Olympus includes that of an afterburner whereas the RFP engine is without an afterburner.Table 6.II - RFP Dimensions Compared to the SNECMA OlympusRFP SNECMAOlympusFan Face Diameter 50 in 47.5 inLength 310 in 280 inWeight 7200 lbs 7000 lbsMax Dry Thrust 26,356 lbs 31,350 lbsBased on these data the RFP engine resembled outdated technology; a more efficient and modern engine will beneeded for the design of the Vendetta. The RFP engine deck was used as a baseline for designing a newer, better engine,as it was the only full engine deck available. It was determined that an F119 engine would be the initial design engine forthe airplane. This engine is currently used in the F-22 and a derivation of the engine (the F135) is to be used in the F-35.Engine performance data for the F119 are classified except that it is in a 35,000 lbs thrust class. Several methodswere utilized to narrow in on the thrust produced by the F119. Through the use of The Integrated High Performance27

Turbine Engine Technology (IHPTET) program, a program through the Air Force Research Laboratories in collaborationwith aerospace companies, F119 characteristics were estimated. IHPTET, which began in 1988 and should culminate in2005, consists of a three-phase plan, utilizing the most current advancements in industry. IHPTET’s goal is to producerevolutionary advancements in turbine engine technologies by combining advanced material developments, innovativestructural designs and improved aerothermodynamics. The three-phases of the program are shown in Table 6.III.Table 6.III - IHPTET Goals+100% Thrust/WeightPhase III (2005)-40% Fuel Burn+60% Thrust/WeightPhase II (1997)-30% Fuel Burn+30% Thrust/WeightPhase I (Completed)-20% Fuel BurnThe Air Force Research Laboratory (AFRL) states that Phase I of the program has been completed and that thetechnology has been applied to existing engines including the F100, F110, F404, and the F119. Based on these dataPhase I of IHPTET was applied to the RFP engine deck to yield an F119 engine, both the RFP engine and the F119 arelow bypass turbofan engines. The 20% decrease in fuel burn was applied and then the weight was decreased by 22.5%and the thrust increased half of a percent to account for the 30% change in thrust-to-weight ratio. The resultinguninstalled thrust produced by the F119 is 26,500 lbs, has an uninstalled cruise TSFC of 0.95 lbm/lbf-hr and a dry weightof 5,500 lbs.The RFP recommends that future advancements and technologies be taken into account. The remaining phases ofthe IHPTET program have yet to be accomplished; therefore, other advancements in turbofan engines were sought out.The Versatile Affordable Advanced Turbofan Engine (VAATE) is an industry projection to 2020. Even though itbuilds upon IHPTET it uses the F119 as a base engine for its future goals. Figure 6.1 illustrates the goals for turbofanengines through 2020 and Phase I goals of a 25% decrease in TSFC and a 45% decrease in cost by 2010.28

Figure 6.1 - VAATE GoalsDepartment of Defense Office of the Deputy Under Secretary of Defense Mr. Paul F. PiscopoIt is likely that this program will face similar problems in achieving its goal by 2010, in which case a decrease of15% in TSFC was taken and an estimated 25% decrease in cost over the F119. The 15% change in TSFC was achievedby increasing the uninstalled thrust by 13% and decreasing the fuel flow 4%. The new VAATE technology engine has asea level uninstalled thrust of 30,000 lbs and an uninstalled cruise TSFC of 0.80 lbm/lbf-hr, however once inlet andducting losses are accounted for the cruise TSFC is 0.90 lbm/lbf-hr. The resulting uninstalled engine deck for Vendetta issupplied in Appendix A. The engine deck is correct to static sea level conditions similar to the engine deck provided bythe RFP.The uninstalled military thrust and TSFC of the engine at various altitudes can be seen in Figure 6.2 and Figure6.3, respectively. The afterburner model was created based on information given in a presentation on the LockheedMartin JSF test program. The approximate afterburner thrust was given as 40,000 lbs, and the dry thrust at 27,000 lbs.This resulted in a maximum thrust 1.5 times military. A maximum TSFC of 2 lbm/lbf-hr was used as most modernengines produce a value around 2 for maximum TSFC.29

35,00030,000Sea Level 1.5K 5K10K20K 25K 30K 36,089Thrust (lbF)25,00020,00015,00010,0005,00043K50K55K60K65K70K00 0.5 1 1.5 2 2.5 3Mach NumberFigure 6.2 - Thrust Curves for Altitudes from Sea Level to 70,000 ft1.2TSFC (1/hr)10.80.60.4Sea Level1.5K5K10K20K 25K 30K 70K65K60K55K50K43K360890.200 0.5 1 1.5 2 2.5 3Mach NumberFigure 6.3 - Military TSFC Curves for Altitudes from Sea Level to 70,000 ftOnce the engine deck had been generated the engine dimensions were once again considered. The fan facediameter of the new engine was assumed to be that of the current General Electric F136 engines that are being tested.They are currently using a fan face diameter of 48 inches (even though production engines will use a 43 inch diameterfan face). Low bypass turbofans typically have smaller fan face diameters; however, as the bypass ratio increases the fanface diameter would increase as well. Since future technology is being taken into account it is likely that the engine thatwould produce this thrust would have a larger bypass ratio but a smaller core keeping the fan face diameter comparableto current sizes. The length of the engine was estimated based on lengths of recent engines. Engines used for comparisoninclude the F100, F101, F110, and F404. Engine lengths varied from about 150 to 200 inches. The engine length wasdetermined to be 192 inches as this engine is a more advanced engine requiring higher thrust production.30

6.2 InletsInlet sizing for supercruise at Mach 1.6 restricted inlet choices to a one, two or three shock inlet. Figure 6.4demonstrates shock relationship to Mach number. A pitot inlet is good up until about 1.6 Mach and it is by far thecheapest inlet possible. However, the performance of the inlet above Mach 1.6 is very poor. The pressure recovery of atwo shock inlet (one oblique and one normal shock) and a three shock inlet were analyzed.Figure 6.4 - Shock Angles for Design Mach Number -Mattingly, Heiser and Daley Aircraft Engine DesignAs can be seen, there is a lower cost associated with a two shock inlet. More shocks drive the inlet to be larger,longer, and send multiple radar returns. The above traits do not show enough of a benefit to go with a three shock inlettherefore a two shock inlet was chosen.The optimum deflection angle for Mach 1.6 flow was found for a two shock inlet by finding the stagnationpressure loss across the oblique and normal shock for different deflection angles. The results were graphed in Figure 6.5and the resulting deflection angle for the greatest pressure recovery was found to be 10.75 degrees yielding a pressurerecovery of 97.65%.31

0.990.980.97Pressure Recovery0.960.950.940.930.92Deflection Angle = 10.75Pressure Recovery = 0.97650.910.90.890 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15Flow Deflection AngleFigure 6.5 - Optimum Deflection Angle for Mach 1.6 FlowFinding the optimum deflection angle for a three shock inlet is more involved; therefore, a rough estimate of a sixdegree deflection angle followed by another 6 degree deflection angle was used to compare against the two shock inlet.The difference in on design pressure recovery is about 1%; however, the larger the deflection angles become the betterthe pressure recovery will become and the longer the inlet. The pressure recovery comparison can be seen in Figure 6.6.The military specification for inlets is given below and is represented in the graph.Mil Spec MIL-E-5008BηrSpec⎧ 1 M ≤ 1⎩1 0.075( 1) 1 50= ⎨ −1.35M0 − < M0

Figure 6.7 shows the off-design inlet area ratio that is required for the Vendetta. The equations used to find thedata are shown below. The actual inlet capture area is depicted by A 1 , with the area at the shock being A s , and the actualflow area being captured by the inlet as A 0i . As the Vendetta climbs, the engine requires a greater amount of inlet area fora constant mass flow rate.Mass Flow RatioA A A=A A A0i 0i s1 s 1Area RatioAA0i s ssρ V=ρ V0 0Inlet Area Ratio1.31.21.110.90.80.7Design Point0.60.50.40.30.20.100 0.5 1 1.5 2 2.5 3Mach NumberFigure 6.7 - Off Design Area Required for Engine Mass FlowThe inlet capture area was found by first estimating the mass flow rate required by the engine at the design point.The mass flow of the engine could be estimated using the following equation.•Mass Flow Estimation ( ) 2em= 26⋅FrontFaceDiameterThe front face diameter of 4 ft was used; this yielded a mass flow rate of approximately 415 slugs/sec. Now usingthe mass flow equation shown below, the area of the inlet could be found for the design mission.Mass Flow Equation•em=ρ AVOnce this was done the mass flow equation was used to calculate the area at different altitudes based onconservation of energy. For the desired design point of 1.6 Mach and an altitude of 55,000 ft it was found to be about 6ft 2 . Since different parts of the mission take place at several different altitudes above 50,000 ft, the inlet area was sized to6.5 ft 2 . By sizing the engine to 6.5 ft 2 air could be bypassed from the inlet to the air cooled fuel cooler. The inlet has a33

oundary layer diverter for high speeds and auxiliary doors for low speed flight, since the required inlet area at take offwill be twice what it is at cruise. The final inlet sizing for Mach 1.6 is:Inlet capture area = 6.5 ft 2Inlet compression angle 10.75 degreesInlet Pressure Recovery is 97.6%Speed after Normal Shock, M=0.82Figure 6.8 - Vendetta S-Duct Side ViewThe inlet is located on a boundary layer diverter on the lower side of the wing. This keeps any vortices producedoff of the wing or side of fuselage from being ingested by the inlet, as well as aid in inlet capture at high angles of attack.6.3 S-DuctS-ducts were used to move the flow from the inlets to the engine faces tohide the compressor face of the engine so it could not be seen. An S-duct frontalview is shown in Figure 6.9. The red parallelogram outlines the inlet while thedotted circle outlines the engine face, as can be seen the engine face cannot beseen thru the inlet. Stealth is a requirement for the mission and the compressorface is a large contributor to radar return.Figure 6.9 - S-Duct Front ViewThe S-duct goes from a minimum area just aft of the inlet to a maximum areaat the compressor face as can be seen in Figure 6.8. The S-duct shape progressivelygoes from a square at the inlet to an oval and then a circle at the engine face. Theportion of the S-duct closest to the fan face is used to straighten and slow the flowbefore it hits the compressor. This is done by having that portion of the duct be fairlylong and gradually diffuse up to the compressor face through an upper and lowerdeflection angle of 3 degrees as shown in Figure 6.10.Figure 6.10 - Diffuser Angleto the Engine FaceThe S-duct is 24 feet in length with an averageheight of 2.7 feet. The efficiency of the S-duct is foundusing the S-duct geometry and Figure 6.11. This yielded alength over diameter of just over 10 and an engine area toinlet area of 2 which yielded a duct efficiency of 91.5%.The resulting overall inlet and ducting efficiency is 89%.Figure 6.11 - S-Duct Efficiency34

6.4 NozzleThe nozzle of the Vendetta incorporates an afterburner and thrustreversers. It utilizes a low signature axisymmetric advanced nozzle, developedunder the IHPTET program, similar to that inFigure 6.12. The advanced nozzle is being used because it has comparableFigure 6.12 - Low-SignatureAxisymmetric Advanced Nozzlesignature to that of a 2-D nozzle, however it weights 50% less, costs 60% less and requires 300 fewer parts. Thrustvectoring is will not be incorporated as the Vendetta is not required to maneuver like a fighter.The nozzle will have thrust reversing capabilities to enable the aircraft to land on an icy runway and stop withinthe required 8,000 ft specified by the RFP. Clam shell style thrust reversers, which reverse 25% of the thrust through a15 degree angle, will be used. The thrust reversers will depart from the upper and lower fuselage just prior to the nozzle,translate back and come together behind the nozzle.35

7 Structural Layout & Material SelectionThe overall layout of the Vendetta’s structure is shown in Figure 7.1 and Foldout 2. The wing structure is similarto that of an F-15 and the material selection is similar to an F-22. The main load path is in the form of a central keel thatruns from between the nozzles and engines to the nose gear attachment point. The weapons bay splits the keel in thecenter of the aircraft. The load is shifted from the keel to the aft weapons bay wall and back into the keel at the forwardend of the weapons bay. A close-up of the weapons bay is also shown in Figure 7.1.Main WingSparsWeapons BayStiffenersFigure 7.1 - Structure Buildup for VendettaThe layout of Vendetta’s inlets and landing gear allow for a continuous structural member, in the form of abulkhead, to carry the aerodynamic loads from each wing directly to the central keel. This approach changes as thebulkheads move away from the main wing load paths. The weapons bay splits the forward wing attachment bulkheads.This occurs well in front of the aerodynamic center of the wing. Just forward of the aerodynamic center is the mainforward load path for the wing. The aft load paths are a ring structure around the engines and inlets. The importantthing to note is that where the primary loads are being distributed, between 25 to 50 percent of the mean aerodynamicchord, the bulkheads are continuous. Because of the thin root selection for the Vendetta, care was taken to ensure thatthe wing could withstand the tremendous loads produced by the 7-g load requirement with a factor of safety of 1.5. It36

was determined that each wing would have to withstand 660,000 lb. Though each individual spar was not sized, it wasdetermined that if the main wing spars were 3 in thick (simple beam cross-section) the wing would be able to withstandthe 7-g maneuver.It is also important to note that the landing gear attach to a bulkhead just forward of the aft closure to the weaponsbay. This locates the airborne and ground laden load paths on top of each other, allowing for some redundancy in thestructure and resulting in a lighter aircraft. Another redundant feature is the aft wing main load path. This bulkhead actsas the main forward engine attachment point. Again, this allows for a minimum of large structural bulkheads and thuscreates a lighter aircraft. The wing attachment points are shown in Figure 7.2.The empennage structure follows the same methodology as the wing attachment structure. The vertical tailsattach to the aft primary carry through of the wing. The aft vertical attach point is the same as the primary load path forthe horizontal tails. The horizontal tail is an area of concern for the Vendetta. The horizontal surfaces are capable ofproducing tremendous forces on the aircraft. At full deflection, the horizontal stabilizers could produce a 10,000 poundforce which would be transmitted through the pivot. It was determined that a 4 inch diameter pivot would be capable oftaking the shear and bending stress produced by this force however the structural rigidity be compromised. The root ofthe horizontal was widened to allow for a 10 inch diameter pivot shaft and increased structural rigidity. The loads takenby the pivot must be transmitted to the keel of the aircraft. There is a ring carry through structure that distributes the loadfrom the pivot point to the central keel. Two secondary bulkheads back up this main bulkhead. The empennagestructure is shown in Figure 7.3.ForwardSecondaryBulkheadsMain GearAttachmentForwardPrimaryBulkheadAft SecondaryBulkheadsAft PrimaryBulkhead &Main EngineAttachmentFigure 7.2 - Wing Attachment Detail37

Vertical AttachmentPointsHorizontal Pivot10” Diameter ShaftHorizontal StructuralLoad PathsFigure 7.3 - Empennage Structural LayoutThe structure of Vendetta was created to adhere to RFP load requirements. A V-n Diagram shown in Figure 7.4 wascreated using the required maximum and minimum g’ limits, and knowing the maximum dynamic pressure the aircraftshould withstand. This diagram shows the load envelope the aircraft can operate in. The diagram also shows thestandard gust lines for 1-g flight. The materials selection for Vendetta was a challenge. Vendetta takes advantage of thebenefits of modern composites while relying on the proven durability of more conventional materials. The materialsselection for different components is shown in Foldout 4.8Max g Limit6Max q4Max LiftGust Lines60 ft/secg's200 ft/sec-2-4-60 ft/secMin g Limit0 500 300 1000 600 1500 900 20001200Equivlent Knots Equivalent Velocity Airspeed(ft/sec)Figure 7.4 - V-n Diagram for Vendetta38

Hydraulic System (Orange)Electrical System (Blue)Fuel System (Red)Materials SelectionID1234567891011121314ActuatorNose RetractFlow Nose Deflector SteeringFlow DeflectorLE FlapMain RetractBrakesMPRL RotateTE FlapAileronEngine Driven PumpAPU Driven PumpHorizontalRudderThrust Reverser499Number1126242444224448151617345,67810 11 1,2ID151617181920ComponentRADARAvionicsSystemsDeIcing BootPTO driven GeneratorAPU driven Generator18ID212223242526Number11221Pneumatic System (Green)Pneumatic SystemAvionics CoolingOBIGGSFlow PressurizationDeflectorLAU-142APHXEngine BleedNumber1112122325212227292432332830ID272829303132333435Forward Fuselage Tank23,034 # JP-831ComponentAAR PortFlow Single Deflector Point RefuelingFlow Forward Deflector FuselageTank & PumpACFCWing Tank & PumpAft FuselageTank & PumpPrimary Fuel PumpEngineAPURight Wing Tank6,366 # JP-8Number111221221444440363738404340423945ID3637383940414243444546474849ComponentRadomeCanopyFlow Forward Deflector Skin & ChineBulkheads / FramesFuel TankWing RibsInlet DuctWeapons Bay DoorsControl Surface SkinLanding GearWing SparsHigh Temperature StructureHorizontal Pivot ShaftHorizontal Internal Structure4146MaterialCompositePolycarbonateFlow Composite Deflector & TitaniumResin transfer, Molded CompositeAluminum, TitaniumResin Transfer, Molded CompositeResin Transfer, Molded CompositeCompositeThermoplastic Skin, Resin TransferMolded Composite StiffnersCompositeTitanium, Steel AlloyTitanium, Resin TransferMolded CompositeTitaniumCarbon CompositeAluminum12132019263435Aft Fuselage Tank23,208 # JP-844474814Total Fuel59,000 # JP-84449Chris DroneyNate SchnaibleRev. 3VendettaScale 1:150Kolby KeiserChris MaglioHigh RollersFoldout 4Systems andMaterial SelectionChris AtkinsonDan Salluce5/23/02

8 Landing GearLanding Gear design for the Vendetta has eight significant design drivers.1) Tire selection to permit a high 150 knot takeoff and landing speed2) 120,000 lb gross weight3) Ease of loading and reloading weapons4) Tail Strike Angle5) Ground Handling Characteristics6) Structural Location7) Minimal Internal Volume Usage8) Low WeightSuitable structural attachment points dictated the main gear be positioned near the subsonic center of pressure onthe main wing (near the main spar) shown in Figure 7.2. This placement, as well as limited internal volume, goodground handling characteristics, minimal frontal area, and ease of unloading and loading weapons led to the adoption ofa tricycle landing gear configuration. The main gear configuration was then approximated as a 737 type main gear, (nearthe Vendetta’s TOGW) for volume purposes.Initial sizing began with tire selection. The main gear of the Vendetta should carry 92% of the TOGW and thenose gear should carry 8%. Starting with a database of tires and wheels the initial listing was narrowed to the choice of36in x 11in tires for the main gear and 24in x 7.7in for the nose gear. The tires selected allowed a 1.5 factor of safety(RFP imposed) over the dynamic landing load of the aircraf.Knowing the approximate volume of the 737 landing gear configuration with usable tires a solid model of thefuselage and internal components was produced todetermine the exact gear location. The initial designconsidered the smallest internal volume as well assmallest frontal area for a given load (Figure 8.1). Afteranalyzing both the internal position the gear would haveto fold into, behind and under the main inlet ducts, thetandem configuration was chosen. The main gear was thendesigned to fold into the allotted space; the retractionscheme can be seen in Figure 8.2The next challenge presented was obtaining theFigure 8.1 - Landing Gear Configuration Trade Studynecessary gear height for easy loading and unloading as40