AP05_4.vp Notation AR Wing aspect ratio. CLmax Maximum Lift coefficient of the aircraft with re- tracted flaps. CLmaxFF Maximum Lift coefficient of the aircraft with full flaps. CLmaxL Maximum Landing Lift coefficient of the aircraft. CLmaxTO Maximum Take Off Lift coefficient of the aircraft. RC, RCmax Maximum Rate of Climb. S Wing area. SLG, STOG Landing Ground run, Take Off Ground run. t Time. tmin Minimum time of climb to altitude z. V(RCmax) Speed at maximum Rate of Climb. Vmax, Vmin Maximum level speed, minimum level speed. Vs, VsFF Stalling speed flaps up, stalling speed flapsdown. WE Empty Weight. WTO Maximum Take Off Weight. P Power. z Altitude. �, �0 Density, density at sea level. 1 Introduction The class of Ultralight (ULM) and light aircraft in general has attracted by growing interest through Europe in recent years. Only in Italy in the last 5–6 years, at least 10 companies have started production of ULM aircraft. There is a very ac- tive market for this class, used to promote flight at all levels and for sports aircraft The maximum flight speed for ULM aircraft has been increased in recent years through the use of more powerful engines (100 hp instead of 64 or 80) and better aerodynamics. It is not surprising that a maximum level speed of about 280 km/h has been reached. Since the weight constraints are very strict, it is important to study ways to improve structural design, safety, flight qualities, aeroelastic behaviour and systems reliability, without raising costs.. Following the experience acquired in our department in designing light and ultralight aircraft, the design of a new composite ULM is being carried out at DPA. The design goals established for this new design were: 1) Short Take-Off and Landing (STOL) aircraft capable of taking off and landing from an uprepared runway within 40 m; 2) almost complete construction in composite material; 3) foldable wing, in order to make the ULM very easy to use, to put on a trailer and to hangar in a normal size garage; 4) wing with a retractable leading edge slat and slotted/fowler flaps; 5) maximum speed around 190–200 km/h at MTOW of 450 kg; 6) good flight and handling qualities, to be safely flown by inexperienced pilots; 7) low cost. 2 Market survey All the analyzed aircraft are ULM (WTO = 450 kg = 4415 N) and equipped with an 80 hp (59.6 kW) engine; most of them are made of aluminium alloy with a high wing configuration, ensuring high stability and easy piloting. None satisfies all the above-mentioned design goals. In fact, the YUMA, the Savannah and the Zenair CH 701 are successful STOL aircraft made of aluminium alloy; however, their de- © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ 73 Czech Technical University in Prague Acta Polytechnica Vol. 45 No. 4/2005 Design of a Low-Cost Easy-to-Fly STOL Ultralight Aircraft in Composite Material D. P. Coiro, A. de Marco, F. Nicolosi, N. Genito, S. Figliolia The paper deals with the design of an aircraft, starting from a market survey, the conceptual design loop and the preliminary choice of di- mensions, and leading to the detailed design of efficient high-lift systems and a low-drag fuselage shape. Technological challenges regarding the design of low-cost systems for flap/slat retraction and a simple wing folding system are highlighted. Aiming at an efficient optimization algorithm, we developed a new integration technique between CAD, aerodynamic and structural numerical calculation. Examples deriving from this new approach are presented. Keywords: STOL, preliminary design. Aircraft M. W.p. WE [N] W W E TO W S TO [N/m2] S [m2] AR VS [km/h] VSFF [km/h] Vmax [km/h] RC [m/s] STOG [m] SLG [m] CLmax CLmaxFF P92 ECHO 80 a h 2757 0.62 334.43 13.20 6.55 71 61 210 5.5 110 100 1.40 1.90 P96 GOLF 80 a l 2757 0.62 361.84 12.20 5.78 71 61 225 4.5 110 100 1.52 2.06 Table 1: Weights, sizes and performances at sea level of the analyzed aircraft (M. – Material: a – aluminium alloy, c – composite; W.p. – Wing position: h – high, l – low.) sign is unattractive, and they have a fixed slat on the leading edge, which reduces maximum cruising speed. The Sky Ar- row 450T and the REMOS G-3, on the contrary, are high cost “non-STOL” aircraft in composite materials, advanced ULM. They can easily by put onto a trailer, due to their removable or foldable wing. The main characteristics of the analyzed air- craft are shown in Table 1. Their main performance charac- teristics in terms of landing run versus maximum level speed at sea level are shown in Fig. 1. 3 Design point The methodology followed during the design process is similar to that reported in [1], but it has been expressly modi- 74 © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ Acta Polytechnica Vol. 45 No. 4/2005 0 20 40 60 80 100 120 140 160 180 0 50 100 150 200 250 300 Vmax [km/h] L a n d in g g r o u n d r u n [m ] P92 ECHO 80 P96 GOLF 80 REMOS G-3 DF 2000 YUMA (STOL) SAVANNAH (STOL) ZENAIR CH 701 (STOL) AMIGO ! SLEPCEV STORCH Mk4 (STOL) SKY ARROW 450T Allegro 2000 SINUS 912 Motoaliante AVIO J-Jabiru EV-97 EURO STAR Model 2001 JET FOX 97 TL 96 Star Goal position for the aircraft to be designed Fig. 1: Landing ground run versus maximum level speed at sea level Aircraft M. W.p. WE [N] W W E TO W S TO [N/m2] S [m2] AR VS [km/h] VSFF [km/h] Vmax [km/h] RC [m/s] STOG [m] SLG [m] CLmax CLmaxFF REMOS G-3 c h 2757 0.62 366.65 12.04 7.98 75 63 220 6.5 80 140 1.38 1.95 DF 2000 a h 2747 0.62 367.88 12.00 8.33 66 56 215 5.5 110 100 1.79 2.48 YUMA (STOL) a h 2766 0.63 328.46 13.44 7.07 55 50 175 6.0 40 55 2.30 2.78 SAVANNAH (STOL) a h 2668 0.60 343.81 12.84 6.28 50 45 160 6.0 50 50 2.91 3.59 ZENAIR CH 701 (STOL) a h 2580 0.58 387.24 11.40 5.90 53 48 153 7.0 50 50 2.92 3.56 AMIGO ! a l 2806 0.64 339.58 13.00 5.24 74 64 250 6.5 80 100 1.31 1.75 SLEPCEV STORCH Mk4 (STOL) a h 2649 0.60 275.91 16.00 6.76 52 46 155 4.5 50 50 2.16 2.76 SKY ARROW 450T c h 2825 0.64 326.76 13.51 6.96 70 61 192 5.1 120 80 1.41 1.86 Allegro 2000 a h 2727 0.62 387.24 11.40 10.23 73 63 220 5.0 150 100 1.54 2.06 SINUS 912 Motoaliante c h 2786 0.63 360.07 12.26 18.28 66 63 220 6.5 88 100 1.75 1.92 AVIO J-Jabiru c h 2649 0.60 474.17 9.31 9.49 74 64 216 6.0 100 160 1.83 2.45 EV-97 EURO STAR Model 2001 a l 2570 0.58 448.63 9.84 6.67 75 65 225 5.5 125 90 1.69 2.25 JET FOX 97 a, c h 2845 0.64 301.95 14.62 6.54 70 60 175 6.0 100 120 1.30 1.77 TL 96 Star a l 2747 0.62 364.83 12.10 6.87 80 63 250 6.0 90 100 1.21 1.94 fied for the ULM category: in particular, new statistical rela- tions between take off ground run STOG and Take Off Pa- rameter for ULM TOPULM (1), landing ground run SLG and landing stall speed VSL, power index Ip (3) and maxi- mum speed at sea level Vmax have been calculated, as shown in Figs. 2, 3 and 4. TOPULM is defined as: TOPULM TO TO L TO � � � � � � � � � � � � � W P W S C� max ; (1) with W S � � � � � � TO in [N/m2] and W P � � � � � � TO in [N/W]; � � � � 0 . (2) Ip is defined as: Ip W S W P � � � � � � � cr 3 (3) with W S in [psf] and W P � � � � � � cr in [lbs/hp]; P kz kv Pcr TO� � . (4) In (4) Pcr and PTO are respectively the power at cruising and take off, kv and kz are the speed and altitude factor (for a four-stroke engine kv � 1 and kz � �1.22), � is the engine admission limit. The data scattering is probably due to lim- ited reliability of the published data, and due to an unbiased difficulty in measuring the data: for example, slight dif- ferences in executed manouvres lead to great differences in measured data. For this STOL aircraft, the main restrictions are maxi- mum speed, take off and landing run, as shown in Fig. 5. Once these limitations have been reported in a graph re- lating power loading (W/P)TO and wing loading (W/S)TO, the resulting shaded area represents all the possible de- sign point choices. Maximum power loading is fixed ( ( W/P)TO � 74 N/kW), because maximum take off weight (450 kg � 4415 N) and power (80 hp � 59.6 kW) have been fixed. In this way only maximum wing loading has been © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ 75 Acta Polytechnica Vol. 45 No. 4/2005 ULM: SLG [m] = 0.038V SL 2 0.641V� SL 0 20 40 60 80 100 120 140 160 180 40 45 50 55 60 65 70 75 VSL [km/h] S L G [m ] FAR23: S LG [m] = 0.02354 V SL 2 Fig. 3: Landing ground run SLG versus landing stall speed at sea level VSL ULM: STOG [m] = 0.0649 TOP ULM 2 + 5.0024 TOP ULM 0 20 40 60 80 100 120 140 160 6 8 10 12 14 16 18 20 TOP S T O G [m ] FAR23: S TOG [m]= 5.22922 TOP 23 + 0.01025 TOP 23 2 Fig. 2: Take off ground run STOG versus take off parameter TOP chosen, based on the criteria for keeping the wing area as small as possible (mainly for cost reasons) and using ap- propriate values of maximum take off and landing lift coef- ficient ((W/S)TO � 324 N/m 2, S � 13.6 m2, CLmaxTO � 2.45, CLmaxL � 3.12). 4 Preliminary design The conceptual loop is shown in Fig. 6. It looks simple, but, for example, converting the geometry of sections into CAD geometry is a complicated and delicate step: aircraft 76 © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ Acta Polytechnica Vol. 45 No. 4/2005 170 180 190 200 210 220 230 0.7 0.75 0.8 0.85 0.9 0.95 Ip V m a x a t s e a le v e l [k m /h ] Fig. 4: Maximum speed Vmax versus power index Ip 64 66 68 70 72 74 76 78 80 82 84 250 260 270 280 290 300 310 320 330 340 350 (W/S)TO [N/m 2 ] (W /P ) T O [N /k W ] Take Off Distance Limit Maximum Cruise Speed Limit Landing Distance Limit Rate of Climb with All Engine Operative Limit Design Point CLmaxL� 3.2 CLmaxTO� 2.6 Chosen power loading (W/P)TO Fig. 5: Maximum power loading (W/P)TO versus maximum wing loading (W/S)TO Geometry of sections Attempt at CAD geometry Section generation, mass and inertial data, surface grids and FEM (Finite Element Method) Semi empirical aerodynamic and structural calculations Performance and flightquality Virtual simulation Have the design goals been achieved? Parametric optimization Detailed calculations, wind tunnel and flight tests yesno wished Complicated and delicate step! Fig. 6: Conceptual loop design surfaces must be carefully defined, otherwise the aircraft geometry will be different from the desired design. The para- metric optimization loop is shown in Fig. 7. First of all, the preliminary geometry was fixed, analyzing existing aircraft and applying semi-empirical methods. The wing was sized to minimize the required power at cruising speed. Some airfoils were analyzed and a new airfoil was designed (modifying NACA GAW1 airfoil) to provide a compromise between lift, drag and pitch moment coefficients. The high lift system and aileron sizing ensures the STOL characteristic and good lat- eral control; this has been demonstrated by J. Roskam [2], W. McCormick [3], C. D. Perkins and R. E. Hage [4] and by the authors [5]. In particular, two possible high lift system configurations are shown in Fig. 8. The horizontal and verti- cal tails were sized by the volume method, ensuring good stability and control also in landing. The fuselage design is very important and it was based on aerodynamic, ergonomic and line of sight studies, as shown in Fig. 9. A 3-view of the aircraft is shown in Fig. 10; Table 2 reports the main dimen- sions, weights and loadings. © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ 77 Czech Technical University in Prague Acta Polytechnica Vol. 45 No. 4/2005 Attempt at geometry defined as n parameters Goal function and constraints definition Ex.: Fusolage drag decrease binding the diameter to a prefixed value (ergonomics) and unmodifying structural endurance First numerical analysis Parameter modification (respecting constraints) according to mathematical “logics” to find goal function optimum (minimum) Numerical analysis (aerodynamic, structural, performance and flight quality) Has the goal function been minimized? Stop no yes Fig. 7: Parametric optimization loop Take off 20°15° Landing 40°15° 40°15° 15°15° (a) (b) Take off 20°15° Landing 40°15° 40°15° 15°15° (a) (b) Fig. 8: Possible high lift system configurations: (a) slat – single slot; (b) slat – fowler Fig. 9: Ergonomics and line of sight of the fuselage Fig. 10: 3-view of the aircraft 5 Numerical analysis The design was accomplished using a code named AEREO [5], which has been developed in recent years at DPA to predict all aerodynamic characteristics in linear and non-linear conditions (high angles of attack) and all flight performances as well as dynamic behavior and flight qualities of propeller driven aircraft. The figures below report some aerodynamic characteristics (Figs. 11, 12, 13 and 14) and per- formance characteristics (Fig. 15) of the aircraft calculated with AEREO code. Table 3 reports the main performances of the aircraft. Further optimization of the global configuration is in progress to improve the wing aero-structural behavior as well as the relative position of the wing and horizontal tail to minimize downwash and induced drag. 78 © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ Acta Polytechnica Vol. 45 No. 4/2005 Czech Technical University in Prague DIMENSIONS, EXTERNAL WING AIRCRAFT Span [m] 9.71 Length overall [m] 6.52 Root chord [m] 1.40 Height overall [m] 1.35 Tip chord [m] 1.40 Aspect ratio 6.93 PROPELLER (fixed-pitch) Incidence [deg] 2.00 Blade number 3 Diameter [m] 1.66 HORIZONTAL TAIL Span [m] 2.80 AREAS Root chord [m] 0.72 Wing [m2] 13.60 Tip chord [m] 0.72 Ailerons [m2] 1.22 Aspect ratio 3.90 Leading edge flap [m2]: slat 2.04 Trailing edge flap [m2]: single slot 2.85 VERTICAL TAIL Horizontal tail [m2] 2.01 Span [m] 1.47 Vertical tail [m2] 1.08 Root chord [m] 0.87 Tip chord [m] 0.61 WEIGHTS AND LOADINGS Aspect ratio 2.00 Empty weight 280 kg 2747 N Incidence [deg] 0.00 Max T-O and landing weight 450 kg 4415 N Leading edge sweep angle [deg] 22.20 Max wing loading 33.09 kg/m2 324 N/m2 Trailing edge sweep angle [deg] 13.00 Max power loading 5.63 kg/hp 74 N/kW Table 2: Main dimensions, weights and loadings PERFORMANCE (Max weight, ISA, at sea level) Max speed [km/h] 194 Take off run to 15 m [m] 121 Cruising speed [km/h] 165 Landing run from 15 m [m] 100 Stall speed [km/h]: flaps up 65 Landing run [m] 50 flaps down: slat – single slot 48 Theoretical ceiling [m] 7908 Max rate of climb [m/s] 6.69 Service ceiling [m] 7317 Take off run [m] 55 Table 3: Performances 6 Conclusion The preliminary design of a STOL ULM aircraft and nu- merical performance prediction has been shown. The aircraft shows acceptable performances that are consistent with the desired design goals. The predicted performances were ob- tained with AEREO code, which confirmed its usefulness as a fast and reliable design tool for propeller-driven aircraft. The parametric design and optimization loops have been high- lighted. Detailed design and optimization of the high-lift sys- tem and three-dimensional aerodynamic analysis are in prog- ress, while wind tunnel tests (high-lift airfoil, aircraft model) are planned in the near future. © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ 79 Czech Technical University in Prague Acta Polytechnica Vol. 45 No. 4/2005 -0.5 0 0.5 1 1.5 2 0 0.05 0.1 0.15 0.2 0.25 CD C L � � �18° � � �15° � � �12° � � �6° � � 0° CDeq � Fig. 11: Polar curves parameterized in d (horizontal all-movable tail deflection) and equilibrium polar curve -1 -0.5 0 0.5 1 1.5 2 -10 -5 0 5 10 15 20 25 �b [deg] C L � = -18° � = -15° � = -12° � = -6° � = 0° CLeq � Fig. 12: Lift coefficient of the aircraft versus alpha body (incidence angle measured in regard to the thrust axis) parameterized in � (hori- zontal all-movable tail deflection) and equilibrium lift coefficient -0.6 -0.4 -0.2 0 0.2 0.4 0.6 -0.5 0 0.5 1 1.5 2 CL C M � � � �� � � � �� � � � �� � � ��� � � ��� � � ��� � � �� � Fig. 13: Pitch moment coefficient versus lift coefficient parameterized in � (horizontal all-movable tail deflection) References [1] Roskam, J.: Part I: Preliminary Sizing of Airplanes. Law- rence, Kansas 66044, U.S.A.: 120 East 9th Street, Suite 2, DARcorporation, 1997. [2] Roskam, J.: Part VI: Preliminary Calculation of Aerody- namic, Thrust and Power Characteristcs. Lawrence, Kansas 66044, U.S.A.: 120 East 9th Street, Suite 2, DARcor- poration, 2000. [3] McCormick, W.: Aerodynamics, Aeronautics and Flight Me- chanics. New York, Chichester, Brisbane, Toronto, Sin- gapore, John Wiley & Sons, 1979. [4] Perkins, C. D., Hage, R. E.: Airplane Performance, Stability and Control. New York, John Wiley & Sons, 1949. [5] Coiro, D. P., Nicolosi, F.: “Aerodynamics, Dynamics and Performance Prediction of Sailplanes and Light Air- craft.” Technical Soaring, Vol. 24, No. 2, April 2000. Prof. D. P. Coiro phone: +39 081 7683322 fax: +39 081 624609 e-mail: coiro@unina.it Dr. A. De Marco e-mail: agodemar@unina.it Dr. F. Nicolosi e-mail: fabrnico@unina.it Dr. N. Genito e-mail: nigenito@unina.it Dr. S. Figliolia e-mail: jdrfig@tin.it Dipartimento di Progetazione Aeronautica (DPA) University of Naples “Federico II” Via Claudio 21 80125 Naples, Italy 80 © Czech Technical University Publishing House http://ctn.cvut.cz/ap/ Acta Polytechnica Vol. 45 No. 4/2005 Czech Technical University in Prague 0 1000 2000 3000 4000 5000 6000 7000 8000 9000 0 5 10 15 20 25 30 35 40 45 50 55 t [min] V[m/s] z [m ] tmin Vmin Vmax V(RCmax) RCmax Theoretical ceiling Service ceiling Fig. 15: Flight envelope -20 -18 -16 -14 -12 -10 -8 -6 -4 -2 0 2 60 80 100 120 140 160 180 200 V [km/h] � e q [d e g ] Fig. 14: Equilibrium horizontal all-movable tail deflection versus speed (center of gravity position is at 25 % of mean aerodynamic chord)