Design Of Ultralight Aircraft - Lisafea
Design of Ultralight AircraftGreece 2018
Main purpose of present studyThe purpose of this study is to design and develop a new aircraft that complies with theEuropean ultra-light aircraft regulations and the US Light Sport Aircraft regulation. For thedesign and development of the aircraft all tools available to the modern engineer have beenproperly used. The aircraft is a two-seater model, oriented towards fast and economictravelling. For this purpose, the development of the wings, the propeller and fuselage has beendone with extra caution, in order for us to achieve the best results possible.The Design ProcessThe procedure below is the one that was followed:1) The airfoil was chosen with the help of xfoil, in order to completely meet therequirements. The results were analyzed by a two-dimensional analysis that wascarried out using the openfoam CFD program.2) Then, the first 3D simulation of the digital model was carried out using the vortexlattice method. At that point, the selection of the design of the aircraft and the aileronwing dimensions for the rudder and the elevator were made, in such a way thateconomy, maximum performance and safe flight are equally achieved. The resultswere checked with OpenFoam.3) With the help of the LISA program, the Finite Element Analysis of the aircraft wasperformed. The wings and the airframe were designed to be able to carry the designloads resulting from the regulations above.4) At this stage, the weight distribution of the aircraft finally became known. The Staticand Dynamic Stability Analysis was carried out with the help of the VLM program.5) By using the OpenFoam program the final and precise analysis of the aircraft’saerodynamics was carried out. The aircraft’s stall behavior was analyzed - atmaximum speed and in all flight combinations- and then the results of the analysiswere evaluated. Additionally, the aircraft’s propeller was also designed.6) A modal analysis was performed in order to calculate the wings’ natural frequencies.With the wings’ aerodynamic data known with the help of LISA, a divergence andcontrol reversal analysis was performed. An unsteady analysis was carried out inOpenFoam in order to calculate the around-the-aircraft unstable load due toturbulence. The results were evaluated in accordance with the natural frequencies thatwere previously calculated with the help of Lisa, and then a flatter test wasperformed.7) With the help of the Code Aster program, an elasto-plastic analysis of the fuselagewas carried out, in the event of a collision.8) The aircraft’s technical characteristics.
1) Airfoil designAs it has already been mentioned, once the aircraft design objective has been established, theprimary topic of study for the engineer is the airfoil. For this specific aircraft, whose goal isdirected towards fast and economic travels at the flight level of 8000-12000 feet, the ideal forthis purpose airfoil was chosen. This airfoil has a particularly low drag when it comes totravel conditions, but if it was to be manufactured in a way that would allow the use ofnegative flaps, it may maintain both a low drag and an ideal buoyant force, even at highspeeds. This is very important, because it is possible for an airfoil to have a low drag, even ata high speed, but to also be able to exert a large buoyant force, which will compel the aircraftto move with a negative pitch in order to maintain its flight level, which will also lead to anincrease of the rest of the aircraft’s drag coefficient, along with a simultaneous increase intravel costs and decrease in maximum speed.Below are figures of the analysis made in FORTRAN environment and the resulting graphs.Figure 1.1: The figure above shows the analysis results in xfoil
Figure 1.2: The figure above shows the results in a graphics environment. Analyses fornegative flap positions were performed, resulting in the creation of an area of constant drag(less than 3 per thousand). At the same time, the value of the buoyancy coefficient varied inorder to guarantee the horizontal motion of the aircraft. This airfoil is ideal for this study’saircraft.
2) 3d vlm aircraft analysisThe airfoil selection was made in the previous section, based on the results of a twodimensional analysis. This happened in order for us to be able to reduce calculation time andto settle on the ideal airfoil, easily and economically.Ιn this section, an analysis of the aircraft as an entity in space will be carried out for the firsttime - in other words, a three-dimensional analysis. When this analysis has been completed(at this stage, many configuration tests will take place in order for us to settle on a design thathas the optimal characteristics), the not-quite-final design of the aircraft will be selected. It isnot quite final yet, because the geometric characteristics may change during the aircraft’sstability testing. Weight distribution has not been finalized just yet and that is why a stabilitytesting cannot be done at this stage of the design. Ιt will be finalized, though, after the finiteelement analysis that follows. The results will also be verified by OpenFoam. It will bechecked whether it meets the study’s requirements, drag in cruise conditions, of maximumspeed and satisfactory buoyance with full-range flaps that will ensure a maximum stall speedin order to meet the criteria of the European Light Aircraft regulation and also to reach highperformance while saving fuel.Figure 2.1: The figure above shows the three-dimensional aircraft with the wing configurationthat was chosen in order to best satisfy the design requirements.
Figure 2.2: The image above depicts the 3D result of the analysis in OpenFoam. Some of theaircraft’s flow lines are also shown, in order for us to understand the horizontal flightaerodynamic performance of the fuselage.At this particular stage we gave the aircraft in its not-quite-final form and the overall plan thatwill allow us to estimate the aircraft’s dimensions with the help of the LISA finite elementprogram is now ready.
3) Finite element analysisAt this stage of the study, all structural parts of the aircraft will be measured using the finiteelement analysis of the LISA program. The design loads are calculated in a way that they aremeeting its category’s requirements of EASA and FAA.The figures below are from the finite element analysis.Figure 3.1: The figure above shows the stress forces exerted due to an 8g load during theflight.
Figure 3.2: The figure above shows the stress forces exerted due to an 8g load during theflight.Figure 3.3: The figure above shows the stress forces exerted due to an 8g load during theflight.It is easily observed that the wings’ maximum expected load (limit load) is 8g.
Figure 3.4: The figure above shows the stress forces exerted due to a 15g hard landing load.Figure 3.5: The figure above shows the stress forces exerted due to a 15g hard landing load.
Figure 3.6: The figure above shows the stress forces exerted due to a 15g hard landing load.Figure 3.7: The figure above shows the stress forces exerted due to an 8g load during theflight.
Figure 3.8: The figure above shows the stress forces exerted due to an 8g load during theflight and a propeller load with a safety factor of 5.Figure 3.9: The figure above shows the stress forces exerted due to an 8g load during theflight and a propeller load with a safety factor of 5.
Figure 3.10: The figure above shows the stress forces exerted due to a 6g load during landing.Figure 3.11: The figure above shows the stress forces exerted due to a 6g load during landing.
Figure 3.12: The figure above shows the stress forces exerted due to a 6g load during landing.Figure 3.13: The figure above shows the stress forces exerted due to a 6g load during landing.
Comments:The wings and fuselage are durable for stress forces exerted due to an 8g load, the fuselagedurable enough to withstand a collision load of 15g. Finally, the engine mounts are durableenough to withstand a propeller load with a safety factor of 5. The landing system has a loadbearing capacity of up to 6g during landing.During collision the fuselage remains within the elastic region up to 15g. It is of a satisfactorysize and so the design of the aircraft can continue. In the next chapter, the behavior of thefuselage frame will be studied with the help of the Code Aster program.
4) Aircraft’s Flight stability analysisThe aircraft’s building materials as well as the method and the cross sections have alreadybeen selected and it is tested that they meet the requirements of the present study. From thisdata the center of gravity and the moments of inertia were calculated. The vlm program wasprogrammed according to these elements, in order for us to perfect the aircraft’s design bycreating a stable and tractable aircraft.Figure 4.1: The aircraft is statically stable and Cm 0 for 0 AΟA. For 0 AOA, Cl 0, theplane is flying. It is noticed that the lift to drag ratio (glide ratio) is very satisfactory, forwhich the very small drag of the aircraft is responsible.
Figure 4.2: longitudinalFigure 4.3: lateral
Comment:The aircraft is also dynamically stable. The center of gravity’s initial estimate was almostidentical to the actual one, yet another finite analysis was done, the aircraft is meeting thedesign goals, so the study can continue.
5) CFD analysis in OpenFoamIn the figures below we see the results from the analysis performed in OpenFoam. In order forus to ensure the safety and performance of the aircraft all possible flight and speedcombinations were studied. From this high-precision analysis the flight program that followswas also created. Also, the characteristics of the propeller (power, speed range, diameter,number of blades and pitch) were selected having taken into consideration the drag data of theanalysis as well as the flight speed.Figure 5.1: Result of the aerodynamic analysis for a flight at maximum speed
Figure 5.2: Result of the aerodynamic analysis for a flight at maximum speedFigure 5.3: Result of the aerodynamic analysis for a flight at approach speed. At this point itwas studied whether the main wing vortices negatively affect the elevator’s performance to anextent that it becomes dangerous for the flight’s safety.
Figure 5.4: Result of the aerodynamic analysis for a flight close to stall speed. At this point itwas studied whether the main wing vortices negatively affect the elevator’s performance to anextent that it becomes dangerous for the flight’s safety. From this angle the loss supportvortexes in the wing root are also visible. Should we have an irrotational flow round theendpoint, the twist (washout) is satisfactory.Figure 5.5: Result of the aerodynamic analysis for a flight close to stall speed. We have anirrotational flow round the endpoint, the twist (washout) is satisfactory.
Figure 5.6: Result of the aerodynamic analysis for a flight close to stall speed. At this pointthe correct performance of the wingtip was studied. It was designed in such way that theairflow produced by the pressure difference between the lower and upper surface of the flapcreates a vortex (known as wingtip vortices), though one that will not hit the top of the flap.This resulted to a higher buoyancy coefficient, lower stall speed and a better behavior as theailerons receive air without vortices.
Figure 5.7: Pressures around the aircraft at cruise speeds.Figure 5.8: Pressures around the aircraft at approach speeds.
Figure 5.9 Pressures around the aircraft at take-off speeds.Figure 5.10: Pressures around the aircraft at final approach speeds.
Figure 5.11: Pressures around the aircraft at speeds just before stall with fully extended flaps.Figure 5.12: Graphs of buoyancy and drag created by the vortex lattice method (jblade)program initially used for the propeller’s design. At this point a two-dimensional analysis ofdifferent airfoils is made in order for us to select the most appropriate combination that willform the propeller flap.
Figure 5.13: Graphs of buoyancy and drag created by the vortex lattice method (jblade)program initially used for the propeller’s design. At this point a 360 degrees analysis of theairfoils is made (for convenience, we only show one). Then, with the help of the Prandtlnumbers they will eventually be shown in a three-dimensional flap.Figure 5.14: Propeller analysis using the vortex lattice method program. We have all thenecessary data to make a choice. After running several tests in the three-dimensional design,the designer concluded that the best propeller was the one with the characteristics above (weshow only the final test and not all of them, for convenience). Moreover, below we see theanalysis results using OpenFoam.
Figure 5.15: Propeller analysis at climb speed with the help of OpenFoam. The figure aboveshows the speeds around the propeller.Figure 5.16: Propeller analysis at climb speed with the help of OpenFoam. The figure aboveshows the speeds around the propeller, as well as the flow lines that indicate the propeller’s“pulling” direction.
Figure 5.17: Propeller analysis at climb speed with the help of OpenFoam. The figure aboveshows the speeds around the propeller, as well as the flow lines that indicate the propeller’s“pulling” direction. The flow is irrotational due to the high performance coefficient of thepropeller. This propeller is indeed ideal for this aircraft.Figure 5.18: Propeller analysis at maximum ground power with the help of OpenFoam. Thefigure above shows the speed profile around the propeller.
Figure 5.19: Propeller analysis at maximum ground power with the help of OpenFoam. Thefigure above shows the vortices around the propeller.Figure 5.20: Propeller analysis at take-off speed with the help of OpenFoam. The figureabove shows the speed profile around the propeller.
Figure 5.21: Propeller analysis at take-off speed with the help of OpenFoam. The figureabove shows the vortices around the propeller.Figure 5.22: Propeller analysis at maximum speed with the help of OpenFoam. The figureabove shows the speed profile around the propeller.
Figure 5.23: Propeller analysis at maximum speed with the help of OpenFoam. The figureabove shows the vortices around the propeller. The flow is irrotational.The results from both software match. The aircraft speed/engine speed diagram is depictedbelow.
Figure 5.24: The diagram above shows the engine speed in relation to the aircraft’s speed inkm/hour as it resulted from the previous analysis. We easily notice the constant speed effectthat was achieved thanks to the meticulous selection of the airfoil and the three-dimensionalset.
6) Static and dynamic aero elasticityA) Static aero elasticityIn an aircraft, two significant static aeroelastic effects may occur. Divergence is aphenomenon in which the elastic twist of the wing suddenly becomes theoretically infinite,typically causing the wing to fail spectacularly. Control reversal is a phenomenon occurringonly in wings with ailerons or other control surfaces, in which these control surfaces reversetheir usual functionality (e.g., the rolling direction associated with a given aileron moment isreversed).i) Divergence occurs when a lifting surface deflects under aerodynamic load so as to increasethe applied load, or move the load so that the twisting effect on the structure is increased. Theincreased load deflects the structure further, which eventually brings the structure to thediverge point. Divergence can be understood as a simple property of the differentialequation(s) governing the wing deflection.ii) Control reversalControl surface reversal is the loss (or reversal) of the expected response of a control surface,due to deformation of the main lifting surface. For simple models (e.g. single aileron on anEuler-Benouilli beam), control reversal speeds can be derived analytically as for torsionaldivergence. Control reversal can be used to aerodynamic advantage, and forms part of theKaman servo-flap rotor design.With the help of the finite element program LISA and the OpenFoam, these two effects weretested and it was found that the aircraft is safe across the entire design speed rate. The wingstiffness is high and it is secured by the two effects above.B) Flutter analysisAn analysis that included the influence of the time variable was performed in OpenFoam. Inthat way the non-steady load due to the aircraft’s turbulence was calculated. The results forthe wing (which are of high importance in the present analysis) are demonstrated below.Then, a modal analysis was made using the LISA program. The results stemming from theOpenFoam analysis were compared to the natural frequencies. There is no flutter at highspeeds. A slight resonance was observed at high speeds, but according to the results fromLISA the wing is able to withstand it. However, the maximum speed allowed was set wellbelow this speed.
Figure 6.1: The figure above shows the results in reference to time when the aircraft flies atflutter speed.The data above were analyzed using the LISA finite element program and after a circularprocess the analysis was completed. The figures below are from the analysis performed inLISA and they also include the aircraft’s flight-envelope diagram.Figure 6.2: The figure above shows the maximum displacement for the 1st eigenvalue
Figure 6.3: The figure above shows the maximum displacement for the 2nd eigenvalueFigure 6.4: The figure above shows the maximum displacement for the 3rd eigenvalue
Figure 6.5: The figure above shows the maximum displacement for the 4th eigenvalueFigure 6.6: The figure above shows the maximum displacement for the 5th eigenvalue
Figure 6.7: The figure above shows the maximum displacement for the 6th eigenvalueFigure 6.8: The figure above shows the maximum displacement for the 7th eigenvalue
Figure 6.9: The figure above shows the maximum displacement for the 8th eigenvalueFigure 6.10: The figure above shows the maximum displacement for the 9th eigenvalue
Figure 6.11: The figure above shows the maximum displacement for the 10th eigenvalueFigure 6.12: The figure above shows the stress forces resulting from a dynamic responseanalysis (for loads in flutter condition) in the LISA finite element program.
Figure 6.13: The figure above shows the displacement resulting from a dynamic responseanalysis (for loads in flutter condition) in the LISA finite element program.Figure 6.14: The figure above shows the speed resulting from a dynamic response analysis(for loads in flutter condition) in the LISA finite element program.
7) Airplane hard landing (as a result of stalling during flotation)Figure 7.01: The figure above shows the stress forces resulting from a non-linear impactanalysis in the Code Aster finite element program. The stalling condition near the groundwas emulated (using results from OpenFoam) and the worst case scenario was chosen (theheight is such that the aircraft will be landed on the runway at a high angle speed but it is notsufficient enough for corrective flotation). The fuselage is strong enough to endure this whileprotecting the life of the passengers, however, in its front part there were areas that thematerial almost reached its strength resulting in extensive delamination damages, thoughwhich was acceptable as it helped absorb the collision energy.
8) Technical characteristics of aircraftFigure 8.1: The figure above shows the thrust or drag in reference to velocity.Figure 8.2: The figure above shows the rate of climb in reference to velocity.
Figure 8.3: The figure above shows the flight envelope diagram of the aircraft.The detailed technical characteristics of the aircraft are shown below, as they resulted fromthe analysis above.ModelClassificationGeneral LayoutAccommodationsUltra-Light AirplaneConventional2 seatsAirworthiness RequirementsAircraft TypeAirframeWing ConfigurationTail ConfigurationPower Plant ConfigurationLanding Gear ConfigurationMultipurposeCompositeLowY-Fuselage mountedSingle-engine, Piston, Tractor, Fuselage mountedFixed, Nose, Fuselage mountedLength Overall6,37 mHeight Overall1,850 mTotal Wetted Area44,888 m²
WINGArea9,900 m²Span9,000 mRoot chord1,300 mTip chord0,900 mTapered ratio1,444Aspect ratio8,182Longitudinal position on the fuselage1,690 mSweep angle0,0 Sweep angle at 25% of wing chord0,0 Sweep angle at 50% of wing chord0,0 Dihedral3,0 Standard mean chord1,100 mMean aerodynamic chord1,120 mWetted area17,617 m²Ratio - Wing area vs Total wetted area0,221Ratio - Wing wetted area vs Fuselage wetted area1,113Ratio - Wing wetted area vs Total wetted area0,392FLAPERONSArea1,525 m²Span (each)3,850 mRelative span (both)85,50 %Standard mean chord0,202 mRelative chord18,00 %Position along the wing span0,650 mLocation along the span14,44 %Hinge axis relative position9,0 %Maximum down deflection40,0 Maximum up deflection-15,0 Ratio - Flaperon span vs Wing span0,855
Ratio - Flaperon area vs Wing area0,154TAILSTails area4,410 m²Tails wetted area8,952 m²Tails area / Wing area0,446Ratio - Tails wetted area vs Total wetted area0,203HORIZONTAL TAILTypeStabilizer and elevatorArea2,810 m²Span2,950 mRoot chord0,950 mTip chord0,950 mTapered ratio1,00Aspect ratio3,56Longitudinal position on the fuselage4,97 mSweep angle at leading edge0,0 Incidence0,0 Relative incidence0,0 Standard mean chord0,950 mMean aerodynamic chord - Chord0,950 mAIRFOIL CHARACTERISTICSAirfoilMaximum relative thicknessLocation of maximum relative thicknessLeading edge radiusLift slope - airfoilAirfoil - zero lift angleNACA 66-0099,1 %45,0 %0,7 %0,104/ -0,1
Lift slope - Tail aloneAerodynamic center positionTail wetted area0,074/ 5,328 m1,666 m²Ratio - Tail area vs Wing area0,085Ratio - Tail area vs vertical tail area0,474Ratio - Tail area vs Total wetted area0,020Ratio - Tail wetted area vs Wing wetted area0,094Ratio - Tail wetted area vs Fuselage wetted area0,105Ratio - Tail wetted area vs Total wetted area0,038ELEVATORArea0,793 m²Span2,480 mRelative span84,0 %Relative chord35,0 %Position along the spanHinge axis position0,177 m10,0 %Maximum down deflection20,0 Maximum up deflection-30,0 Ratio - Elevator span vs Horizontal tail span0,840Ratio - Elevator area vs Horizontal tail area0,294VERTICAL TAILTypeFin and rudderArea1,600 m²Span1,600 mRoot chord0,700 mTip chord1,300 mTapered ratio1,86Aspect ratio3,20Longitudinal position on the fuselage4,870 m
Root to tip sweep23,20 Standard mean chord1,000 mMean aerodynamic chord - Chord1,030 mTail moment arm3,239 mAIRFOIL CHARACTERISTICSAirfoilMaximum relative thicknessLocation of maximum relative thicknessLeading edge radiusLift slope - airfoilAirfoil - zero lift angleLift slope - tail aloneTail wetted areaNACA 66-0099,1 %45,0 %0,7 %0,104/ -0,1 0,046/ 3,248 m²Ratio - Tail area vs Wing area0,161Ratio - Tail area vs Horizontal tail area0,626Ratio - Tail area vs Total wetted area0,035Ratio - Tail wetted area vs Wing wetted area0,184Ratio - Tail wetted area vs Fuselage wetted area0,204Ratio - Tail wetted area vs Total wetted area0,074RUDDERSpan1,520 mRelative span95,0 %Relative chord40,0 %Position along the spanHinge axis position0,070 m50,0 %Maximum left deflection35,0 Maximum right deflection-35,0 Ratio - Rudder span vs Vertical tail span0,950
FUSELAGELength5,870 mMaximum height1,110 mMaximum Width1,120 mLength of constant section0,000 mFuselage frontal form coefficient0,960Fuselage lateral form coefficient1,773Fuselage frontal areaWetted area1,001 m²15,833 m²BASEBase frontal form coefficient0,960LANDING GEARBaseMaximum tail down angleWetted area1,519 m8,0 3,243 m²MAIN GEARFixed gearMain gear - Tire6.00-6Main gear - Tire diameter445 mmMain gear - Tire width160 mmAUXILIARY GEARRetractable gearAuxiliary gear - Tire5.00-5Auxiliary gear - Tire diameter361 mmAuxiliary gear - Tire width126 mmENGINEEngine number1
Engine modelSubaru EA-71Engine - Specific fuel consumption0,310 kg/kW.hEngine - Specific weight1,10 kg/kWMaximum engine power62,517 kW 85,0 hpMaximum engine rpmPower-to-wing area ratio5750 t/min5,62 kW/m²Power-to-weight ratio0,139 kW/kgWeight-to-power ratio (Power loading)7,198 kg/kWPROPELLERNumber of propellerTypeMaterialNumber of blades1Fixed pitchWood2Propeller pitch angle - Minimum16,0 Propeller pitch angle - Maximum46,0 Propeller diameterDisc area1,700 m2,269 m²Maximum disc loading27,55 kW/m²Maximum disc loading vs Number of blades13,76 kW/m²Spinner - Diameter0,200 mSpinner - Length0,210 mMOMENT OF INERTIA (ESTIMATED)Fuel system - Main tank locationWingFuel system - LocationWingFuel system - Capacity20.lFuel system - LocationWingFuel system - Capacity20.lFuel system - Maximum fuel capacity40.lWing tank capacity40.l
WEIGHT AND LOADINGMaximum Takeoff weight450,0 kgEmpty weight253,9 kgFlight weight450,0 kgUseful weight196,1 kgWeight of crew - Unit86,0 kgWeight of crew - Total172,0 kgWeight of freight - Unit5,0 kgWeight of freight - Total10,0 kgWeight of fuel33,5 kgWeight of crew - Minimum50,0 kgWeight of fuel - Minimum10,0 kgMinimum Takeoff weight313,9 kgPower plant65,0 kgEngines(1)65,0 kgPropellers(1)4,0 kgCOMPUTED WEIGHTWingHorizontal tail65 kg12,2 kgVertical tail12,5 kgFuselage45 kgMain landing gear8 kgAuxiliary landing gear5 kgEngines(1)Propellers(1)65,0 kg4 kgFuel system5,3 kgControl system8,9 kgElectrical systemInstruments10,0 kg3,0 kg
FurnishingsEmpty weight10,0 kg253,9 kgCENTRE OF GRAVITY POSITIONOccupant(1)2,020 mOccupant(2)2,020 mFreight2,570 mFuel1,960 mBatteries (M)1,030 mWing2,250 mHorizontal tail5,300 mVertical tail5,660 mFuselage2,340 mMain landing gear2,540 mAuxiliary landing gear1,910 m(1)Engine0,770 m(1)Propeller0,250 mFuel system1,820 mControl system2,080 mElectrical system1,450 mInstruments1,280 mFurnishings1,930 mFlight weight450,0 kgMASS CORRECTION FACTORGeneral1,000MISSION SEGMENT WEIGHT FRACTION[1] Warm-up1,000[2] Taxi1,000[3] Takeoff1,000
[4] Climb0,997[5] Cruise0,906[6] Descent1,000[7] Loiter1,000[8] Descent1,000[9] Landing1,000[10] Taxi1,000WEIGHT RATIORatio - Empty weight vs Maximum Takeoff weight0,564Ratio - Useful weight vs Maximum Takeoff weight0,436Ratio - Fuel weight vs Maximum Takeoff weight0,074Ratio - Useful weight vs Empty weight0,772Ratio - Fuel weight vs Empty weight0,132Ratio - Fuel weight vs Useful weight0,171Ratio - Weight of engine vs Empty weight0,256Ratio - Empty weight vs Wing area25,647 kg/m²Ratio - Maximum Takeoff weight vs Wing area45,455 kg/m²Ratio - Empty Weight vs Total wetted areaRatio - Maximum Takeoff Weight vs Total wetted area6,530 kg/m²11,574 kg/m²AERODYNAMICSMaximum lift coefficient (Dirty)2.35Maximum lift coefficient (Clean)1,55Maximum lift increment0,80Wing loading at maximum Takeoff weight45,455 kg/m²Wing loading at empty weight25,647 kg/m²Friction coefficient, Coefficient (power flight)Friction coefficient, Reference altitude0,005300.m
QUALITY CRITERIAFuel consumption (cruise)6,27 l/100kmFLIGHT AT MAX CONTINUOUS SPEEDFlight speed245 km/h- Ground speed (GS)245 km/h- True Air Speed (TAS)245 km/h- Indicated Air Speed (IAS)218 km/hAirplane CG rel. position (%CMA)28,00 %Wing loading45,455 kg/m²Flight weight450,0 kgFlight altitude2400.mRange384 kmEndurance1 h 33 minTime to climb7 min 52 sPower, maximum62,157 kWPower, available62,000 kWPower, required60,000 kWEngine relative powerSpecific fuel consumption96,5 %0,310 kg/kW.hEngine rpm5500 t/minPropeller - rpm2750 t/minPropeller - Pitch angle24,25 Propeller - Efficiency85,0 %Propeller - Thrust (net)1489 NRATE OF CLIMBMAXIMUM RATE OF CLIMBFlight weight450,0 kgFlight altitude0.m
Rate of climb6,1 m/sFlight speed165 km/h- Ground speed (GS)165 km/h- True Air Speed (TAS)165 km/h- Indicated Air Speed (IAS)165 km/hPower, maximum62,157 kWPower, available62,000 kWPropeller - rpm2400 t/minPropeller - Pitch angle24,25 Propeller - Efficiency79,47 %Propeller - Thrust (net)1095 NPropeller - Thrust-to-Power ratio17,66 N/kWClimb angle7,58 Climb slope13,43 %TAKEOFFAirplane CG rel. position (%CMA)Runway surface28,0 %ConcreteTakeoff run185.mTakeoff distance to 15m284.mTakeoff weightFlight altitude450,0 kg0.mWing trailing edge deflection angle10,0 Runway slope0,0 %Front wind speed0 km/hAt rotation speedStall speedTakeoff speed71,5 km/h120 km/hLift coefficient (maximum)1,88Lift coefficient0,68
Mean accelerationRunway surface2,96 m/s²grassTakeoff run236.mTakeoff distance to 15m335.mTakeoff weightFlight altitude450,0 kg0.mWing trailing edge deflection angle10,0 Runway slope0,0 %Front wind speed0 km/hLANDINGAirplane CG rel. position (%CMA)28 %Runway surfaceConcreteLanding weight450,0 kgFlight altitude0.mWing trailing edge deflection angle40,0 Runway slope0,0 %Front wind speed0 km/hBreakdownSpeed, approach125 km/hSpeed, flare out102 km/hSpeed, touch down95 km/hLanding, brakes OFFDistance from the obstacle (15m)565.mDistance during approach90.mDistance during flare out35.mDistance during touch down40.mDistance during ground roll400.mMean deceleration0,844 m/s²Landing, brakes ONDistance from the obstacle (15m)260.m
Distance during approach90.mDistance during flare out35.mDistance during touch down40.mDistance during ground roll95.mMean deceleration3,76 m/s²BEST RANGERange915 kmFlight altitude2400.mFlight speed182 km/h- Ground speed (GS)182 km/h- True Air Speed (TAS)182 km/h- Indicated Air Speed (IAS)162 km/hAirplane CG rel. position (%CMA)28 %Flight speed (optimal) (104,4 kg/m²)182 km/hEndurance5 h 2 minFlight weight450,0 kgWing loading45,455 kg/m²Wing loading (optimal) (182 km/h)45,455 kg/m²Power, maximum62,157 kWPower, available62,000 kWPower, required22,000 kWEngine relative powerSpecific fuel consumptionPropeller - rpm35,5 %0,300 kg/kW40,467.h1950 t/minPropeller - Pitch angle24,25 Propeller - Efficiency79,55 %
Figure 8.4: 3d view of the aircraft.
STABILITYLONGITUDINAL DERIVATIVESLATERAL DERIVATIVES
AcknowledgmentI’d like to thank LISA’s technical supportI also want to th
Design of Ultralight Aircraft Greece 2018 . Main purpose of present study The purpose of this study is to design and develop a new aircraft that complies with the European ultra-light aircraft regulations and the US Light Sport Aircraft regulation. For the
the 3DES key, the MIFARE Ultralight C design shows some differences compared to MIFARE Ultralight design. In order to minimize the impact of chip differences for the customer designs and ease the transition from existing MIFARE Ultralight systems to MIFARE Ultralight C systems, this document highlights the differences between the two products.
The maximum mass is lower than or equal to 300 kg for a single-seat aircraft and 450kg for two-seat aircraft. Class 2 (Ultralight trike): An ultralight trike is a motorized card suspended from rigid sail. It is effectively a powered hang glider. Class 3 (Airplane): An Ultralight airplane is an aircraft featuring a fixed wing. In France, ULM
Frame kit Ultralight Frame kit Ultralight Ultralight Evo Ultralight C14 C13 XS 470 115 530 402 72,2º 74,8º 517 382 SM 480 120 543 402 72,5º 74º 523 383 MD 505 150 557 402 73º 73º 549 384 LA 540 185 570 402 73º 72,5º 583 386 XL 552 195 583,3 402 73,5º 72,5º 594 396 XS 470 115 530 402 72,2º 74,8º 517 382 SM 480 120 543 402 72,5º 74º .
Keywords: ultralight aircraft, aerodynamics, composites, shell construction, glass fibre, epoxy resin, engine block carrier. Abstract The goal of this research was to make airworthy, to modify, and to improve the performance of a Weedhopper II ultralight aircraft (ULA) (JC-24)) manufactured in 1986. In this paper, the renovation of
to assist amateur and ultralight builders in developing individualized aircraft flight-test plans. It also provides guidance for experimental light sport aircraft flight testing after modifications to the aircraft. It provides recommendations and suggestions you can combine with other sources on
MIFARE Ultralight EV1 - Contactless ticket IC Rev. 3.3 — 9 April 2019 Product data sheet 234533 COMPANY PUBLIC 1 General description NXP Semiconductors developed the MIFARE Ultralight EV1 MF0ULx1 for use in a contactless smart ticket, smart card or token in combination with a Proximity Coupling Device (PCD).
- B734 aircraft model added - B735 aircraft model added - E145 aircraft model added - B737 aircraft model added - AT45 aircraft model added - B762 aircraft model added - B743 aircraft model added - Removal of several existing OPF and APF files due to the change of ICAO aircraft designators according to RD3: A330, A340, BA46, DC9, MD80
Fedrico Chesani Introduction to Description Logic(s) Some considerations A Description Language DL Extending DL Description Logics Description Logics and SW A simple logic: DL Concept-forming operators Sentences Semantics Entailment Sentences d 1: d 2 Concept d 1 is equivalent to concept d 2, i.e. the individuals that satisfy d 1 are precisely those that satisfy d 2 Example: PhDStudent .
