The Elite: A high speed, low-cost general aviation aircraft for Aeroworld

coefficient at fifty stations along the wing. The stall model incorporated compared these section C e's to the C emax of the airfoil. The airfoil was assumed to have a linear slope until its Cemax was reached and then to linearly drop off at the same rate if its C fmax was exceeded. In essence, this put the following Cg vs. alpha curve for the airfoil into the lifting code.

Figure 4.4.1: Airfoil Model Incorporated into the Lifting Line Code C_ max C_ same slope alpha With this airfoil model, the code searched for the condition where the total integrated wing CL was the greatest. This should give a more accurate approximation of when the wing will stall. The integration of the wing CL distribution was found to be a maximum at 12.3 degrees. The wing CL distribution at this condition is shown below.

Figure 4.4.2: Wing C L Distribution at Stall

1.2

"T'IF'W-- 0.8 ._] o c- O 0.6- _j II I!

O.

0 0.5 1 1.5 2 2.5 3 3.5 4 Spanwise Location, (ft) Aerodynamically, the larger the aspect ratio the better. CLaw increases and the induced drag decreases as the aspect ratio increases. Unfortunately, this makes the wing long and narrow which is inherently structurally less sturdy. After discussions with the structures group, it was decided that the largest feasible aspect ratio was 9. Thus the wing aspect ratio was set at 91 Below is a summary of the of the main wing aerodynamic characteristics. These characteristics resulted from the decisions made above based on the desire to increase the cruise speed while still maintaining adequate low speed performance.

1For the airfoil selection process the aspect ratio was assumed to be 8. However, the airfoils should perform the same relative to one another regardless of the aspect ratio.

Table 4.4.1: Summary of Wing Characteristics 4.76/rad CLaw CLo 0.0995 CLmax 1.088 e 0.978 O_stall 12.33 ° Recruise 347,000 Retakeoff 133,000 Aspect Ratio 9 Planform Area 6.5 ft 2 Planform Shar_e Rectangle / 7' 7 3/4" Span Cord 103/16" Airfoil DF101 4.5 Empennage Planform and Airfoil Selection For the tail section, we examined the follow four airfoils: NACA 0009, NACA 0012, SD8020, and a flat plate. The SD8020 was chosen for the horizontal tail and a flat plate for the vertical tail. The SD8020 had the best response in the zero angle of attack regime and had the best drag characteristics. However, one of our main design goals was the aesthetics of the airplane so it was decided to sweep the tail surfaces. It appeared to be a very difficult task to build the vertical tail in the desired trapezoidal shape with an airfoil cross-section so the design group settled upon a fiat plate geometry. This difficulty was avoided in the horizontal tail by making the planform shape a parallelogram. Thus an airfoil cross-section could be used in the horizontal tail. The actual sizes of the tail surfaces were determined by the stability and control analysis.

4.6 Drag Breakdown Since a high cruise speed is an essential part of the DR&O, drag is an extremely important issue. Unfortunately, the geometry of this aircraft was fairly fixed by other concerns (payload size, placement of servos, landing criteria etc.) so not much could be done to affect the drag. The Elite does incorporate some cosmetic changes to help with the drag but their actual impact in a low Reynold's number regime is difficult to quantify.

The main cosmetic change over past designs is to have a very smooth, sleek fuselage.

This helps to reduce drag while at the same time enhancing the aesthetic nature of the aircraft.

For the drag breakdown, two different sources were used to obtain a value for CDo. The first was Dr. R. C. Nelson's (ref. 4) breakdown method presented in AE441.

This breakdown method uses empirically determined CDo values for each component.

The second column of CDo'S came from a variety of data sources. In the second column, the wing profile drag is assumed to be the same as the DF101 airfoil. The fuselage drag came from a fuselage drag chart on page 180 in ref 5 with a fineness ratio of 7.6. The vertical tail drag is assumed to be that of a flat plate. The horizontal tail has the CDo of the SD8020 airfoil. Finally, the landing gear is assumed to have the CDo of a right circular cylinder. Once the individual CDo'S are obtained for each part of the aircraft, the basic equation for the total drag coefficient is: Below is a tabular listing of the CDo values used in determining the drag breakdown. The planform area of the wing was used as the Srefio_ = 6.5 ft 2 Table 4.6.1: Drag Breakdown Components for Two Different Sets of Data Component Sref Sref (ft 2) CDo-Nelson CDo-Data Planform 6.5 0.007 Wing 0.008 Cross Section 0.136 0.11 Fuselage 0.075 Nacelles Cross Section 0.049 0.06 0.06 Vertical Tail Tail Area 0.55 0.008 0.013 Horizontal Tail Tail Area 1.8 0.008 0.008 Frontal Area 0.0958 0.95 I. 1 Landing Gear Interference +10% +10% CDo Result 0.0292 0.0325 From this table, one can see that the landing gear is the major contributor in terms of the drag. In fact, the landing gear accounts for 50% of the total CDo. One solution to this problem would have been the use of retractable landing gear. Unfortunately, the impact of this option was not fully realized in the initial concept selection process. In the future, the use of retractable gear should definitely be considered. The drag polar is given by: C D = CD. + CD, C[ where, CD, - 1 rtA Re which for the Elite equals: CD= 0.0325+ 0.03974C_.

This relationis plottedbelow.

Figure 4.6.1: Aircraft Drag Polar

1 1.2 CL One thing to nonce is the profile drag is very large compared to the induced drag.

Any uncertainty in the profile drag estimates can cause dramatic changes in the performance projections for the aircraft. For The Elite, the pessimistic drag estimates were used to compute the performance to help insure that the objectives were met. Once a relationship between CL and CD has been established, the Lift/Drag curves for the airplane are easily obtained. The L/D is plotted against the angle of attack and level flight velocity below.

Figure 4.6.2: L/D Curves -- a) vs. Angle of Attack b) vs. Level Flight Velocity 14.

15- ° 12.

1 1o-

° 10.

/

8- £3

s /

\

£3 6-:

o ¢

\

4-

/

2-

-s- /

0- -10_;/ -6-4 -2 0 2 4 6 8 101214 Velocity, ft/sec Alpha, deg a) b) One thing to notice about the L/D plots is that L/Dmax occurs near stall rather than at cruise. It is desirable to have the L/Dmax occur at cruise since that is the condition were one would fly for the maximum endurance. In the case of The Elite, the maximum range speed is 25 ft/sec a far cry from its cruise speed of 60 ft/sec. The main reason for L/Dmax occurring near stall is the shallow nature of the drag polar. The large value for CDo with a relatively normal value for CDi never allows the drag to increase at a faster rate than the lift. Thus the L/D maximum occurs near the maximum lift condition.

4.7 Aircraft Aerodynamic Summary These values Below is a summary of the aircraft aerodynamic characteristics.

completely describe the aerodynamic quality of the Elite aircraft.

Figure 4.7.1: CL vs (_ of Aircraft Table 4.7.1: Summary of Aircraft Aerodynamics 1.2- 1" 4.76/rad CLct 0.8 CDo 0.0325 0.6 CDi 0.03974 0.4 e 0.86 ..J 0.2 0 L/Dmax 14

,.2"

0-: L/Dcruise 5

.#

-0.2 Planform Area 6.5 ft 2 -0.4 Aspect Ratio -6 -4-2 0 2 4 6 8 10 12 14 Angle of Attack, deg 4.8 References 4.1 Selig, Michael S., Donovan, John F., and Fraser, David B. Airfoils at Low Speeds. Virginia Beach: H.A. Stokely, 1989.

4.2 Miley, S. T., A Catalog of Low Reynolds Number Airfoil Data for Wind Turbine Applications, Texas A&M, 1982 4.3 Fay, Jonathan P., Lifting Line Code written and validated for AE360 under the direction of Dr. E. J. Jumper, Department of Aerospace and Mechanical Engineering, University of Notre Dame, Spring 4.4 Nelson, R. C., "Subsonic Drag Estimation: Component Build-up Method." Department of Aerospace and Mechanical Engineering, University of Notre Dame, 1993 4.5 McCormick, Barnes W. Aerodynamics, Aeronautics, and Flight Mechanics. New York: John Wiley and Sons, 1979.

4.6 Lissaman, P.B.S. "Low Reynolds-Number Airfoils", Annual Review of Fluid Mechanics, Vol 15, 1983, pg 223-239

PROPULSION SYSTEM DESIGN DETAIL

5.1

Requirements and Objectives Design Requirements: 1. Environmentally safe.

2. High speed performance.

3. Maximum take-off distance of 28 feet on smooth runways and 42 feet on rough runways.

4. Propulsion system installation under 20 minutes.

5. Ability to fly to nearest alternative airport and loiter for one minute.

6. Fuel stored in the wing carry-through structure.

Design Objectives: 1. Minimum Cruise Speed of 60 feet/sec.

2. Maximum velocity of at least 80 ft/sec.

3. Short and Rough Field Take-Off Capability 4. Range of 30,500 feet to allow service of all Aeroworld airports 5. Aesthetics 5.2 System Selection The design of the propulsion system involved the selection of the motor, propeller and fuel system. In order to be environmentally safe to Aeroworld, the RFP required that the propulsion system of the aircraft employ a state-of-the-art electric propulsion system (Ref. 5.1). The driving factors that determined the components of the propulsion system were the maximum and cruise velocities, the aircraft range, take-off distance, cost and aircraft maximum take-off weight. The final system consisted of the Astro Cobalt 15 engine, the Zingali 10-8 three-blade propeller, and thirteen Panasonic P-130SCR battery cells. The following is a table of the aircraft's values used in the calculations.

Table 5.2.1: Aircraft Data Aircraft 4.88 lbs Weight 6.5 ft 2 Wing Area Aspect Ratio

CDo 0.0325

CLmax 1.03 0.89 Oswald Efficiency 5.3 Motor Selection Two motors were considered for The Elite, the Astro 15 and the Aslxo 25. A third

motor,the Astro 05, was available but not considered due to its insufficient power output

leading to an inability to satisfy the take-off distance objective. The motor weight and cost, maximum velocity attainable, and the take-off distance drove the motor selection.

Table 5.3.1 illustrates the advantages, particularly in maximum velocity and weight, the Astro 15 motor has over the Astro 25 motor.

Table 5.3.1: Comparison of Astro 15 motor with Astro 25 motor 28 oz 38 oz Motor Weight Motor Cost $107 $174 71.7 ft/sec 56.5 ft/sec Maximum Velocity 23.6 ft 40+ ft Take-Off Distance (28 ft max) Recommended Motor RPM 16,500RPM 10,000RPM The maximum velocity and the take-off distance comparison were all calculated using our selected propeller, the Zingali 10-8 three-blade propeller, and the battery specifications of a pack voltage of 15.6 volts and a battery capacity of 1300 mah. These predictions were obtained using the FORTRAN programs TAKEOFF (Ref. 5.2), PROP123 (Ref. 5.3), and PAVAIL (Ref. 5.4).

Intuitively, the larger more powerful Astro 25 should have out performed the Astro 15. This would have been the case if the propeller was large enough to take advantage of Astro 25's extra torque capabilities. With a smaller ten inch diameter propeller, however, high RPM's are more important than high torque capabilities. The Astro 15 motor was ultimately selected because of its lower weight and superior velocity performance attributable to the motor's much higher maximum RPM. In addition, the Astro 15 cost significantly less than the Astro 25. Table 5.3.2 contains further information about the Astro 15 motor characteristics.

Table 5.3.2: Other Astro 15 Characteristics Name Astro Cobalt 15 Maximum Power 185 watts Internal Resistance 0.12W Gear Ratio 31:14 95% Gear Efficiently kt 1.0978 in-oz/amp Tloss 1.3729 in-oz kv 7.8568 10 -4 V/rpm The gear efficiency was assumed to be 95% based on recommendation from AE454 Propulsion class (Ref. 5.5).The motor torque and battery constants are taken from the curve fit of the Astro 15 motor performance based upon motor data supplied by the manufacturer (Ref. 5.6). The plots used to determine the motor torque and battery constants are shown in Figure 5.3.1.

Figure 5.3.1: Motor Torque and Battery Constants - 20000 _"_._ Kv=7.s56S_V/_rln j -18000

.,,- 16000

b_ O - 14000 O

.,,, .120o0

15.

Kt=l.097_ " -10000 o O o [- -"_m r -8000 10- _..6000 O

.t _4ooo

/'- Tout = 1.0978 ia-1.3729 _2000 Nm = -285.428 "m+ 19270A2 F .... I .... I .... I .... I .... I .... [ 0 0 5 10 15 20 25 30 Motor Current (a) It should be noted that the PAVAIL program did not take into account a gear efficiency and torque losses (Tloss). This was compensated by writing our own program to

determineall theperformance characteristics of the propulsionsystem.The programwas

validatedby settingthe gearefficiencyandtorquelossto 100%andzero,respectively,

andcompared to PAVAIL

5.4 Propeller Design Propeller selection proved critical in the aircraft's ability to fulfill the requirements for maximum velocity and take-off distance outlined in the DR&O. Several parameters were examined during the propeller selection including diameter, pitch, manufacturer, and the number of blades. The geometric chord and thickness versus the blade radius where recorded and inputted into PROP123 to attain the results. The program accounted for induced velocity and tip losses, and Reynolds and Mach number corrections. A trade study was performed to determine the effects of the propeller diameter, pitch, manufacturer and number of blades on propeller efficiency, thrust coefficient and power coefficient. For aesthetic purposes, propeller diameter was limited to 10 inches and 11 inches so that the propeller would be somewhat proportional to the aircraft. In addition, a smaller diameter would require shorter landing gear, and thus help reduce weight and drag penalties. The propellers that were considered were the two- bladed Top Flight 10x6, Zinger 10x7, and Zinger 1 lx7; and the three-bladed Zingali 10x8 and Graupner 1 lx7. Figure 5.4.1 illustrates the improvement in thrust coefficient three- blade propellers have over two-blade propellers.

Figure 5.4.1: Thrust Coefficient as a Function of Propeller 0.16 ED i; • Top Flight 10x6 0.14 • Zinger 10x7 o "_ 0.12 -- "1 !

• Zinger 1 lx7 ._ • oo : 0.1 -• a °:a [] Zingali 10x8 o i• • 1 o Graupner 11x7 = =, ie o *

° I '+° I I

i • o 0.06 m ' 0.04 i I 0.02 .... , .... u.... t .... , .... ,'..., .... , .... I ....

0.3 0.35 0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 Advanco Ratio

Figure5.4.2illustratesthatalthougha three-blade propellerhasimprovedthrust

coefficient,it alsohasa higherpowercoefficientthantwo-bladepropellers.

Figure 5.4.2: Power Coefficient as a Function of Propeller • Top Flight 10x6 • Zinger 10x7 • Zinger 1lx7 * Zingali 10x8 [] Graupner 1 lx7 O n Advance Ratio Figure 5.4.3 compares the efficiencies of three-bladed propellers and two-bladed propellers. As one can see, only the Zingali 10x8 had comparably high effmiencies with the best two-bladed propellers.

Figure 5.4.3: Propeller Efficiency as a Function of Propeller 0.85 ¢ _1,.,.,oF,maanlj Rmi, B,,_1 !

• Top Flight 10x6 0.8 t • Zinger 10x7 I • • = 0.75 • Zinger 11x7 nO_ 0 0 0 A m 0.7 [] Zingali 10x8 s ¢ o Graupner 1 lx7 _0.65 ¢ ,iL s no 0.-6 i i

° 1

0.55 , .,.I. ,,.i .., ,i., .,I....ii.,,I,.,.i ,.., m 0.3 0.35 0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 Advance Ratio Three-bladed propellers where chosen for their higher thrust and power capability, and comparably high efficiencies. Another reason for the choice of a three-blade propeller was the more aesthetic aerodynamic appearance over two-bladed propellers. Aesthetics, again, was one of the major drivers of the design.

The selection of three-bladed propellers was very limited due the fact that few companies manufactured propellers in the desired pitch and diameter. Two three-bladed propellers were acquired and analyzed. These propellers were the Zingali 10-8 and the Graupner 11-7. Figure 5.4.4 shows the advantage the Zingali propeller had over the Graupner propeller in maximum velocity and power available at higher velocities.

Figure 5.4.4: Propeller Maximum Power Available verses Velocity 10-8 lille • Graupner 11-7 20.

0.

2O 30 40 50 60 70 80 Velocity fit/s) Table 5.4.1 is a break down of the propeller performances. The values are taken from a program written to determine power available, power required, range, endurance and current draw at different velocities (Ref. 5.7). Again, the Zingali's performance is superior to that of the Graupner in every category except at maximum rate of climb.

Table 5.4.1: Propeller Performance Comparison Zin[ali 10-8 Graugner 11-7 71.7 ft/s 68.0 ft/s Maximum Velocit 7 80% 71% Cruise Propeller Efficiency Current Draw at Cruise 10.2 amp 10.8 amps Maximum Rate of Climb 15.2 ft/s 15.36 ft/s 27401 ft 25997 ft Ranse at Cruise Endurance at Cruise 7.6 minutes 7.2 minutes The Zingali 10-8 was ultimately selected for two reasons. Lower diameter requires a smaller landing gear which thus decreases the weight and decreases the drag while satisfying the requirement of a rough field capability. A drag breakdown performed by

theAerodynamics Groupindicatedthatthe landinggearaccounted for nearly50%of the

totalparasitedrag,CDo. In addition,theZingali 10-8hada highermaximumvelocity

which wasoneof theprimarydriversof the design.

Figure5.4.5containsthe Zingali's average propellerefficiencyasa functionof

advance ratio. This figure indicatedthatatcruise,the Zingali propellerwasoperating

very closelyto the propeller'smaximumefficiency.

Figure 5.4.5: Zingali Propeller Efficiency as a Function of Advance Ratio 0.82 0.8 >, 0.78 •_ 0.76 ¢.9 _-= tu 0.74 c 0J a_ 0.72 Q_ o a. 0.7

¢,/

0.68 Minimum Operating _ _umn Operating Advance Ratio Advance R_io F 0.66 0.35 0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 0.8 0.85 0.9 Advance Ratio Due to some uncertainty that arose in the PROP123 program, there was considerable concern in the accuracy of its predictions, particularly at low and high advance ratios.

However, as one can see, the propulsion system operates in a relatively narrow band in the linear region of the curves. Thus, within the operating advance ratios, performance predications were expected to be relatively accurate. Nonetheless, past wind tunnel data indicated that the propeller efficiencies found for the Zingali were too high. It was expected that the propeller efficiency of the fiberglass Zingali propeller would be higher than that of wooden propellers because of the capability of machine precision manufacturing, but the improvement was so dramatic as to cause suspicion of the results.

The original PROP123 was used because it provided the best reasonable values of efficiencies to wind tunnel data.

5.5 Engine Control & Battery Selection

Speed and rate of climb was controlled by varying the throttle. During take off, the throttle should be opened fully to a voltage of 15.6 volts and then reduced until the required velocity is attained. The same would be done for climb maneuvers with the throttle varying between the cruise throttle and the maximum throttle. Figure 5.5.1 illustrates the power required and the power available for The Elite at several throttle settings during various flight regimes.

Figure 5.5.1: Effect of Throttle Setting on Power Available 140 I n Oo O00 O0 On o n [] 12.6 V _ Velocity o £x, 60' 4O 2O 0 ,

I ! i

i I I i I I | | i | | i I | l I | I 2O 50 60 70 80 Velocity (ft/s) Table 5.5.1 summarizes estimates of the throttle settings that will attain the desired flight conditions for The Elite.

Table 5.5.1: Throttle Setting for Desired Flight Condition Throttle il Veloci_ Flight Condition 24.8 ft/s 42% Take-off Velocity/Stall 36.9 ft/s 52% Maximum Range at WMTO 49.6 ft/s 65% Maximum Velocity for Range Goal of 30,500 ft 60.0 ft/s 81% Cruise Velocity 71.7 ft./s 100% Maximum Velocit'), The Elite was designed to be powered by 1.2 volt rechargeable battery cells. The batteries chosen for The Elite were the P-130SCR 1.2 Volt batteries having a rated capacity of 1300 mah. This batteries were selected for the technical demonstrator based on results from PAVAIL which again proved to be erroneous when gear efficiency and torque losses were not taken into account. Figure 5.5.2 illustrates the range and endurance for The Elite when gear efficiency and torque losses axe taken into account.

Figure 5.5.2: Range and Endurance Versus Velocity at WMTO 34000 21 33000-

\

32000- i • 31000- e-.

E 30000- o tj o 29000- 13

• i

28000- ° _ 11 e- ............. • o,.

27000- .9 26000- .7 250O0- 24000 , 5 l I I i _ i I I [ I I I I l ] I I I l I a l l 2O 3O 40 50 60 70 8O Velocity (ft/s) As one can see, the range requirement at cruise was not attainable with this battery. In order to have a range of 30,500 feet at a cruise speed of 60 ft/s, 1500 mah would be

required. However,suchbatteries werenot available.The closesmah-rated batteryare

the 1400mahbatteries.Thoughall thecalculations in this reportarebased on the 1300

mah batteries, it is suggested that the 1400 mah batteries should be used for the actual production of The Elite. For performance, only range and endurance would be effected.

With the 1400 mah battery, the maximum range was 35,500 feet at 37 ft/s, maximum velocity for goal range was 57 ft/s, and the range at cruise was 29400 feet. This is very close to our object performance goals.

Nonetheless, with 1300 mah batteries, The Elite can handle 95% of the possible flight routes (see Performance Section). Thirteen cells were chosen based upon the manufacturer's suggested battery pack voltage for the Astro 15 motor. Table 5.5.2 contains the specifications on the 1300 mah batteries.

Table 5.5.2: Battery Specifications 13 Cell Pack Panasonic P- 130SCR 11 Cell Battery 15.6 V Voltage 1.2 V 1300 mah Capacity 1300 mah 78 mW Internal impedance 6 mW 22.1 oz Weight 1.7 oz.

Cost $4.00 $52.00 5.6 Installation One of the requirements of the propulsion system was that it could be installed and removed from The Elite in under 20 minutes. To achieve this, the batteries would be sealed together and placed in the wing carry-through structure. The wing will be able to be screwed off the bottom of the fuselage allowing easy access to the batteries and radio control equipment. The batteries will be fixed within the wing box spars with velcro.

The motor will slide into the nose mount attached to the firewall with four mounting screws. A nose cone and spindle will be mounted for aerodynamics and aesthetics. Further detail on engine and battery mount structure can be found in the Structures section of this document.

5.7 Propulsion System and Performance Summary Motor Astro Cobalt 15 Propeller Zingali 10-8 13 Panasonic P-130SCR Battery Tekin Model Speed Controller Futaba 4N-BL/Attack Radio Control System 2.04 lbs Weight Cost $172.58 Performance 71.7 ft/s Maximum Velocity 60.0 ft/s Cruise Velocity 49.57 ft/s Maximum-Range Velocity 24,099 ft Maximum-Velocity Range 27,401 ft Cruise Range 32,919 ft Maximum Range 5.6 minutes Maximum Velocity Endurance Cruise Endurance 7.62 minutes 14.9 minutes Maximum Range Endurance 5.8 References 5.1 Batill, Dr. Stephen. "High Speed, Low-Cost General Aviation Aircraft for 'Aeroworld.'" Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

5.2 Batill, Dr. Stephen. TAKEOFF FORTRAN program. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

5.3 Batill, Dr. Stephen. PROP123 FORTRAN program. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

5.4 Batill, Dr. Stephen. PAVAIL FORTRAN program. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

5.5 Dunn, Dr. P.F. "AE454 - Propulsion Class Notes - Electric Motor - Propeller Propulsion (with application to Remotely Piloted Vehicles)." Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1993.

5.6 Batill, Dr. Stephen. Group A Design Databook. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

5.7 Le, Tuan. A FORTRAN program to compute range and endurance, 1994.

6 WEIGHT ESTIMATE DETAIL

6.1 Level Zero Weight Estimate A preliminary component weight breakdown is presented in Table 6.1.1. These estimates were based on the data base of prior airplane designs in Aeroworld. The initial weight estimate was a low value of 4.2 lbs. Several of the components were taken directly from RPV catalogues. These include the motor, servos, receiver, speed controller, propeller, and batteries. The wing, fuselage, and empennage weights were all estimated as 2/3 to 3/4 of the values observed in past airplane designs. An uncertainty of + 10 % was added to find the high and low end weight estimations.

Table 6.1.1: Zero Level Weight Component Breakdown Component Vb_ ight IWeight % Structure 0.75 Wing 17.9 Empennage 0.1t 3.8 Fuselage 0.5 11.9 Landing Gear 0.35 8.3 Subtotal 1.76 41.9 Control Systems Servos 0.113 2.7 Receiver 0.059 1.4 Speed control 0.11 2.6 System batteries 0.125 3.0 Subtotal 0.407 9.7 Propulsion 0.64 Motor (Astro 15 w/gear box) 15.2 0.044 1.0 Propeller Batteries 0.94 22.3 Subtotal 1.624 38.7 Payload 0.05 1.1 Total 3.8 + .38 4.2 High-end weight A preliminary center of gravity estimation was made by placing the aircraft components at desirable positions. The position was located at 13.25 inches behind the nose of the airplane.

6.2 Improved Weight and C.G. Estimate A more detailed weight estimation was then calculated after a better understanding of the aircraft layout was obtained. The weight component breakdown is presented in Table 6.1.2. The table is an inclusive summary of all the of the structural weight needed to design the airplane. Each spar that is needed for manufacturing is included. Also in Table 6.1.2 is the x location of the center of gravity of each component measured from the nose of the aircraft and the moment that each creates about the leading edge of the aircraft (the nose). The center of gravity of the entire aircraft resulted in a location of 13.565 inches behind the nose. This was very close to the initial c.g. estimate of 13.25 inches.

Table 6.1.1: Improved Weight Component and C.G. Breakdown System [Part Name X-Location Moment IWeight about nose Propulsion Propeller 0.097 0.000 0.000 Asn'o 15 motor w/mount 0.640 2.500 1.600 Motor Batteries 1.380 12.000 16.560 Speed Controller 0.117 11.500 1.346 Wires batteries--> Speed Controller 0.032 11.750 0.376 motor --> Speed Controller 0.032 7.000 0.224 Propulsion Total Weight 2.298 Avionics iServos Aileron 0.038i 18.000 0.675 Rudder 0.0381 19.250 0.732 Elevator 0.038 19.250 0.732 Receiver 0.059 12.250 0.723 Servo Batteries 0.125 11.500 1.438 Receiver Wires 0.020 15.000 0.300 Antenna 0.010! 25.000 0.250 Control Rods/Links/Horns Rudder Balsa Rod 0.007 26.625 0.197 Control Horn 0.004 34.000 0.150 0.008 Wire Rod with Link (@ 21.250 0.178 servo) 0.008 32.000 Wire Rod with Link (@ 0.268 horn) Elevator BalsaRod 0.007 26.625 0.197 t I Conwol Horn 0.004 34.000 0.150 Wire Rod with Link (@ 0.008 21.250 0.178 servo) Wire Rod with Link (@ 0.008 32.0130 0.268 horn) Aileron Wires 0.037 19.000 0.703 Links 0.004 18.5001 0.065 Nose Gear Sheathed Plastic Rod 0.016 11.625 0.187 0.002 Link (@ gear) 5.000 0.009 0.002 Link (@ servo) 18.250 0.032 0.445 Total Avionics Weight Landing Gear Nose Gear with Horn and Screws 0.150 4.125 0.619 Main Gear 0.256 16.100 4.122 Main Gear Straps and Screws 0.053 16.100 0.850 0.459 Total Landing Gear Weight Payload 0.046 14.500 0.671 Main Wing Leading Edge Location 10.500 Spars 0.1341 Main Top (Balsa) 13.050 0.535 Main Bottom (Balsa) 0.027 13.050 0.352 Leading Edge (Balsa) 0.031 10.560 0.327 Trailing Edge (Balsa) 19.130 0.027 0.517 Secondary Top (Spruce) 16.100 0.017 0.274 Secondary Bottom (Spruce) 16.100 0.017 0.274 Monokote 0.167 15.5901 2.604 Ribs 0.093 15.340 1.427 Aileron 0.032 20.020 0.641 Tips (Soft Balsa Blocks) 0.050 13.750 0.688 Gear Blocks (Spruce) 0.062 15.750 0.983 Fuselage Mating Blocks (Spruce) 0.040 14.575 0.583 Webbing (Balsa) 0.015 12.930 0.194 Hinges 0.053 19.190 1.013 Glue 0.063 18.660 1.166 Fiberglass Spar Mating 0.188 18.660 3.508 0.923 Total Wing Weight Fuselage Longerons Too (Balsa) 0.018 19.000 0.334 Bottom (Balsa) 0.019 19.000 0.361 Port Side 0.009 19.000 0.171 Starboard Side 0.009 19.000 0.171 0.018 Shaping (All 4 Combined) 19.000 0.342 Bulkhead (#'s start @ Firewall) 1 (Spruce) 0.008 4.000 0.031 2 (Balsa) 0.004 6.000 0.027 3 (Balsa) 0.006 10.500 0.066 4 (Spruce) 0.014 12.925 0.178 5 (Spruce) 0.014 13.175 0.183 6 (Spruce) 0.014 15.975 0.224 7 (Spruce) 0.014 16.225 0.227 8 (Balsa) 0.008 20.500 0.167 9 (Balsa) 0.008 24.000 0.192 10 (Balsa) 0.006 28.000 0.168 11 (Balsa) 0.005 32.000 0.160 12 (Spruce) 0.008 34.000 0.272 Monokote 0.050 19.000 0.950 Servo Tray 0.039 15.500 0.605 Tail Cone 0.025 36.000 0.900 0.030 4.000 0.120 Engine Mounting Blocks (Spruce) 0.040 14.575 0.583 Fuselage Mating Blocks (Spa'uce) Glue 0.063 19.000 1.197 0.025 34.000 0.850 Empennage Mating Blocks (Spruce) 0.454 Total Fuselage Weighl Vertical Tail Spars Leading Edge (Balsa) 0.008 34.000 _ 0.286 Hinge Line (Balsa) 0.008 36.000 _ 0.288 Truss Pieces 0.015 35.000 J 0.536 Rudder 0.018 37.0001 0.681 0.001 36.000 0.035 Tip Spar 0.003 34.500 0.099 Root Spar 0.003 29.000j 0.087i Fuselage Blending Block 0.004 36.000: 0.144 Hinges Monokote 0.022 34.000 0.7481 Horizontal Tail Spars 0.009 34.000 0.289 Main Top (Balsa) 0.009 34.000 0.306 Main Bottom (Balsa) 0.009 32.500 0.293 Leading Edge (Balsa) 0.004 39.000 0.156 Trailing Edge (Balsa) 0.034 34,000 1.156 Graphite Rod Ribs 0.024 35.000 0.851 0.020 34,000 0.680 Tip Blocks (Soft Balsa) 0.025 34.000 0.850 Hardwood Connecting Blocks 0.010 Monokote 35.000 0.350 0.030 34.000 1.020 Glue (Vertical and Horizontal) Total Empennage Weight 0257 SUMS ......... > 4.881 66.211 Xcg 13.565 5.125 With 5% Fudge Factor % Cmac of Xcg 0.303 Figure 6.2.1: Weight & Balance Diagram

o

0 2

7 STABILITY AND CONTROL

7.1 Requirements and Objectives Requirements: 1. The stabilator must be able to rotate the airplane at take-off, trim the airplane at cruise and at landing (stall angle), and maintain static stability while in the air.

2. Must execute a steady, level 60-ft-radius turn at 25 ft/sec in order to maneuver in Aeroworld.

3. Must achieve a coordinated turn with the rudder and aileron deflections. The rudder must be able to overcome the adverse yaw created by the ailerons while at a bank angle.

Objectives: 1. Longitudinal static stability must be achieved with static margin > 10%.

This is to enable a novice pilot to fly the airplane.

2. Aircraft is to fly at cruise at zero angle of attack with a zero stabilator deflection in order to minimize the drag of the fuselage.

3. The horizontal and vertical tails and their corresponding control control surfaces should be selected as small as possible, as long as they satisfy stability parameters. This is to keep with the design objective of a lightweight aircraft.

Note: All of the equations an methods used for the stability analysis in this airplane design are taken from (Ref. 7.1).

7.2 Longitudinal Stability The Elite utilized an all-movable tail in order to achieve longitudinal stability and control. This was chosen since it can have a smaller area and still provide the same control power as a conventional tail with an elevator. It also adheres to the design objective of an aesthetically pleasing airplane since its size is proportional to the short fuselage. Past Aeroworld designs had relatively large empennages which looked somewhat awkward on their aircrafts.

The sizing and positioning of the stabilator were initially driven by the need to rotate the airplane at take-off. The stabilator surface area and location were both varied in order to find when they provide a zero moment about the wheel at a certain tail deflection angle. The computer code for this trade study can be found in Appendix C.

The location of the wheel was placed as far forward as possible in order to give a large moment arm for the tail. There was a certain restriction on this position since it had to be a minimum distance away from the main fuselage spar so that loads can be distributed when mounting the wing. The resulting graph is presented below in Figure 7.2.1 Figure 7.2.1: L t vs. Tail Area for Rotation at Take-off

1.t

Tail Deflection Angle 1.4 ° 8 degrees o- 1.2- _9 10 degrees t_L, _ "-- de"gnl _'_''" _--_ 12 degrees <

• 0.8

14 degrees }- 0.6 -43- 16 degrees t- O N "= 0.4- .... o ......... 18 degrees O I 20 degrees 0.2 O.

15 16 17 18 19 20 21 22 23 24 Lt (inches) Given a certain horizontal tail location (It), Figure 7.2.1 shows the tail area and tail deflection angle that was required in order for the airplane to rotate at take-off. It was decided to set the maximum tail deflection angle to + 12 degrees, because this was just below the airfoil's stall angle. Also, greater deflections would interfere with the movement of the rudder as well as create possible problems with the avionics necessary to control the tail. Figure 7.2.1 shows that there was a wide range of acceptable combinations. Therefore, additional measures of merit were needed in order to help determine the tail parameters.

The pitching moment curve slope (Cma) measures the longitudinal static stability of the airplane. When an airplane experiences an increase in its angle of attack due to a positive (nose-up) moment, it must be able to create a negative (nose-down) pitching moment which tends to rotate the airplane back to its equilibrium position. In order for this to occur, the Cma slope must be negative: dCm < 0 dcx The wing, fuselage, and horizontal tail all contribute to the pitching moment of the aircraft. The contribution of the fuselage to the pitching moment in previous Aeroworld airplanes was relatively small. An example of a typical general aviation airplane in (Ref 7.1) also had a negligible fuselage pitching moment. Therefore, since the method for calculating the Cmag of the fuselage is tedious and time-consuming, its effect was neglected. This is one aspect of the design that can be improved upon. An accurate estimate of the fuselage pitching moment would help insure the stability of the airplane since the fuselage contributes a destabilizing effect. The wing also produces a destabilizing effect on the aircraft. Therefore, the horizontal tail of The Elite must provide a large enough pitching moment to overcome the destabilizing effect of the wing and fuselage.

Static margin is the other measure of merit which was explored. The static margin helps to measure the responsiveness of the airplane. It is defined as the distance between the neutral point (XNp) and the airplane's center of gravity position (Xcg), both referenced from the wing's leading edge: Static Margin = XNP X_ C C The neutral point is the furthest aft location that the center of gravity can be located. A center of gravity beyond the neutral point results in an statistically unstable aircraft.

The static margin is normally between 5 and 10 percent for conventional aircraft.

However, for this class of airplanes a slightly larger static margin is desired. This is due to the fact that the pilot is stationed on the ground with limited visual cues resulting in a slower response time than if he were sitting within the aircraft. A larger static margin would make the airplane respond more slowly to control inputs. Thus, a static margin greater than 10% would be desirable for this type of airplane. An experienced RPV pilot may be able to handle an aircraft with a lower static margin, however this airplane was designed for the novice pilot to fly.

The stabilator should then be able to provide a suitable Cma value and a static margin of at least 10 percent while also being able to rotate at take-off. A computer code was written to observe the effects of the tail size and location on the pitching moment and static margin values. The resulting graph can be seen in Figure 7.2.2.

A typical general aviation aircraft (Ref 7.1) has a Cma of -0.68 rad -1. From Figure 7.2.2, the smallest tail area and shortest location possible were chosen which would still give a static margin of at least 10 % and a sufficient Cma value. A smaller tail will provide less weight in raw materials and help with reducing the drag. It also satisfies the design objective of an aesthetically pleasing airplane. A significantly large tail will make the airplane look awkward and dissuade potential buyers. As long as the airplane maintains sufficient handling qualities, a smaller tail should not present a problem.

Because of thesereasons, a tail areaof 0.9 ft 2 and a location of 21 inches behind the

center of gravity was chosen. This provides the airplane with a Cma value of -0.485 rad- 1 and a static margin of 10.5 %. The Cma value was comparable with (Ref 7.1) and other past Aeroworld aircraft. The all-moveable tail was given a sweep angle so that it could maintain its aesthetically pleasing appearance. The angle was arbitrarily chosen to be 15 degrees.

Figure 7.2.2: Variation of Static Margin and Cma 0.4 U 0.35

o 0

iL'

t- | O..

t_ E O .-1 0.05 •-1.5 o D -0.05 .-2 0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 Horizontal Tail Area (sq. ft) A Cm(It=19) • SM (It=19) O Cm(It=20) • SM(It=20) ._4 Cm(It=21) A SM(It=21) [] Cm(It=22) _l, SM{It=22) (9 Cm(It=23) [] SM(It=23) Cm(It=24) O SM(It=24) The following is a table summarizing the design values for longitudinal stability: Table 7.2.1: Longitudinal Stability Parameters 0.9 ft 2 Horizontal Tall Area (S H) SH/S 0.14 VH 0.30 _ mean chord, c 6 inches 21.6 inches span, b AR tail 3.6 Moment arm (lt) 21 inches Tail incidence (it) -2.0 degrees 10.5 % Static Margin XNp 40.5 % Cma -0.485 rad -1 The tail and wing incidence angles were selected in order to enable the airplane to be trimmed at cruise with a zero wing angle of attack and a zero tail deflection. This would make the drag of the airplane at cruise as small as possible since it eliminates any unnecessary drag due to the fuselage at angles of attack. The incident angles necessary to accomplish this were, iw = 0.85 degrees and it = -2.0 degrees. The pitching moment of the airplane as a function of the angle of attack is presented in Figure 7.2.3 below.

It was evident from Figure 7.2.3 that the wing pitching moment curve had a positive slope, and was therefore unstable. The Cmo of the airplane must be a positive value in order for the airplane to trim at positive angles of attack. The Cmo of the wing had a negative value. Therefore, the tall must have a large enough Cmo to counteract the wing's effect. The results can be seen in Figure 7.2.3.

Figure 7.2.3: Pitching Moment Coefficient vs. al 6,_

0.1

0.05 Aircraft

-%

Tail E -0.05 Wing -0.1 -0.15 -0.2 0 2 4 6 8 10 12 14 16 18 20 Angle of Attack (degrees) The Elite, was designed to fly at cruise with a zero angle of attack of the fuselage.

The drag on the fuselage is least when the fuselage was at 0 ° angle of attack. Mounting the wing at the incidence angle equal to the angle of attack needed for cruise thereby minimizes the aircraft's drag. Placing the wing at a particular incidence is a difficult task because of all the imprecision involved. However, for The Elite, the incidence angle was small (less than 1°), so mounting the wing should not pose a problem. For this configuration, the pitching moment should be zero (the aircraft is trimmed) at an a of zero degrees. Figure 7.2.3 shows that this is indeed the case. The equation of the pitching moment curve is : Cm = 0.001 - .009*a where a is in degrees.

The movement of the center of gravity is very small when the passengers and payload are removed. They only make up 1.0 % of the weight and thus do no affect the c.g. location to any extreme. The resulting shift in the center of gravity position is presented in Figure 7.2.4. The two curves are very similar. Even at the forward c.g.

position, the airplane will be able to trim at cruise at essentially a zero angle of attack. As long as the c.g stays in fxont of the neutral point the airplane will remain statistically stable and will be able to trim at positive angles of attack. This requirement did not seem to poseanyproblems.

Figure 7.2.4: Pitching Moment Coefficient for Aft and Forward C.G.

0.05 Cm(aft c.g.)

Cm(fwd c.g.)

-0.05 O E cD -0.1 -0.15 -0.2 -5 0 5 10 15 20 Angle of Attack (degrees) 7.3 Longitudinal Control: The control mechanism of The Elite was the same as the horizontal tail since it utilized an all-moveable tail. It has already been observed that the tail provides enough control in order to rotate the airplane at take-off. It remains to be seen whether the stabilator will be able to trim the airplane at various flight conditions while in the air and at landing. Figure 7.3.1 shows the different trim condition of the aircraft at angles of attack from 0 to 20 degrees. The airplane will stall at an angle of attack just over 12 degrees.

Figure 7.3.1:Effect of Tail Deflectionon Pitching Moment

Tail Deflection Angle .5 -.

dele = -20 deg 0.4! _ _ _ _=.

0.3 '_ dele = -15 deg dele = -10 deg 0.1-! i'_ rA'j _I_ =_"_ "4b-__.,4 _b,,_

I

dele = -5 deg dele = 0 deg -0.2 4 "O' __1.1 h_lhl4 l .... O ..... dele = 5 deg -0.3 -!_ _' -0.4 • -.A--- dele = 10 deg

-o.5

0 2 4 6 8 10 12 14 16 18 20 Angle of Attack (deg) Figure 7.3.1 shows that The Efite could be trimmed at a wide range of angles of attack with minimal tail deflection angles. This enabled the airplane to remain at equilibrium while in flight. When the airplane is landing, it will have an extreme angle of attack close to the stall angle of 12.3 degrees. The stabilator must be able to provide a sufficient pitching moment to trim the airplane with a maximum deflection angle of 12 degrees. Deflection greater than this value were undesirable because the tail will stall at a slightly higher angle. Since it was an all moveable tail, it will have significant control power and a deflection of 12 degrees should be sufficient to trim. The procedure for determining longitudinal control (Ref. 7.1) was simplified since there was an all- moveable tail. For this case, the effective elevator area was the same as the horizontal tail area. The flap effectiveness parameter (t) therefore would simply be set equal to one.

Setting, the a of the airplane at 12.3 degrees and the Cm equal to zero, the corresponding d of the tail could be determined. The tail must deflect an angle of d = -6.2 degrees in order to trim at landing. This was well within its range of + 12 degrees and can thus trim at landing. Similar results can be obtained directly by using Figure 7.3.1.

In Figure 7.3.1, the effect of the stabilator deflection angle on the pitching moment curve was observed. As the tail was given a positive deflection angle (deflected leading-edge up), the pitching moment curve shifted downward. This was because a positive deflection would cause a pitch-down moment on the aircraft. A negative deflection caused the opposite to be true.

Table 7.3.1: Characteristics of Stabilator de (landing) -6.2 de_rees -.218 rad -1 Cmde Se / St 1.0 + 12 ° dem_ 7.4 Directional and Lateral Stability: Directional stability was necessary in order to return the airplane to an equilibrium condition when subjected to a form of yawing disturbance such as sideslip. The requirement for directional stability was for Cnb > 0. The contribution of the fuselage and the wing to the directional stability was determined by way of the equations and graph in (Ref. 7.1). The fuselage and wing create a destabilizing effect on the directional stability. The tail position (Iv) was set equal to the horizontal tall location (lt). The tail area was then varied over a range from 0.2 to 1.0 ft 2 and the different Cnb values were observed. An area of 0.3 ft 2 was chosen which gives a vertical tail volume ratio of 0.011 and a Cnb of 0.025. This is a compatible Cnb value when compared to past Aeroworld airplanes. This area was chosen also based on rudder requirements which will be discussed in section 7.6. There were many uncertainties when calculating the contribution of the fuselage and the wing. They may in actuality contribute more of a destabilizing effect. Therefore, a large vertical tail was used in order to ensure directional stability. The sizing of the vertical tail was re-affirmed when determining the rudder size in the section on directional control.

The lateral or roll stability of the airplane was what enabled it to create a restoring moment when disturbed from a wings-level attitude. For roll stability, the coefficient of the roll moment due to the sideslip should be less than zero (Clb < 0). Since The Elite used ailerons for the turning of the airplane, dihedral was needed only to insure the lateral stability and not turn the aircraft. Therefore, a dihedral angle of 5 degrees was chosen for this airplane design. This angle gives a Clb of -0.104. A table of the aircraft stability coefficients are shown in Table 7.4.1.

Table 7.4.1: Directional and Lateral Stability Parameters 0.3 ft 2 Sv Vv 0.012 S v / S 0.046 mean chord 6 inches Span 7.2 inches ARv 1.2 Cnb 0.025 rad -1 Wing Dihedral G 5.0 degrees -0.104 rad -1 c_ 7.5 Lateral Control Lateral Control would be achieved by the deflection of ailerons. The size and location of the ailerons were determined using the steady state roll equations (Ref. 7.1, Eq. 5.2): i_ = Lpp + L_. _i, CIp The aileron control power is a function of the span, chord, and location of the ailerons (Ref. 7.1, Eq. 2.97): 2CL,,,,, X Y, Ch" - Sb fcydy Y, Note that these equations were included in this report to make clear the method used in determining the roll control of the airplane. The velocity was taken as 28 feet/second to meet the requirement of making the turn in under 30 feet/second and to ensure some margin above the stall speed of 25 feet/second. The maximum aileron deflection angle was set at 15 degrees. The flap effectiveness parameter, t, was found for aileron chord lengths of 1, 1.25, 1.5, 1.75 and 2 inches. At each of these chord lengths, the roll rate was computed for various aileron spans and wing locations.

It was found that a 0.25 inch increase in the chord length would give approximately a 10% increase in the roll rate produced for a given span and location.

Also, a 0.5 feet increase in the span of the aileron for a given chord length increases the roll rate by 25%-30%. (Ref.7.2) suggested that the best way to increase aileron control is

to increasethe aileronspanratherthanincrease the aileronchordbecause increasingthe

chordincreases the effectiveness only slightly but greatlyincreases the loadsplacedon

the servo.Theseadditionalloadsmay causethecontrolrodsto bendor theaircraft's

structureto beslightly distorted. Also thechordshouldnot be so smallthatthe aileron

effectiveness is lost. Thusthe aileronchordlengthwassetat 1.5inches.

A plot of the roll rateversusdifferentaileronspans andlocationswith achord

lengthof 1.5inchesis foundin Figure7.5.1. Notethat theroll ratewasa maximumwith

theinboardedgeof theaileronclosestto thefuselageandwith theoutboardedgeclosest

to thewing tips. However,a roll rate of 40 or 50 degrees/second is not desirablebecause

this wouldbe too fast a roll for this typeof aircraft. Basedon the aircraft'sload factor

while turning,anda 50 feetturningradiusata velocityof 28 ft/sec,the expected bank

anglewasbetween25 and29 degrees. To achievethis bankanglein 1.5seconds, the

requiredroll rate was 16-20degrees/second. Thusit wasdesiredthattheailerons

producea roll rate of atleast20 degrees/second.

Figure 7.5.1: Inboard and Outboard Location vs. Roll Rate Outboard Location 6O of Aileron (feet) 5O

t 4

© L_ 3.5 40.

o o 30.

"""--,

2o

o 0 0.5 1 1.5 2 2.5 3 Inboard Location of Aileron (feet) The manufacturing team suggested that the ailerons run as close inboard as possible so that the control wire would not have to be run very far out along the wing.

Also, (Ref. 7.3) suggested that the ailerons be placed two-thirds to three-fourths of the

way out alongthewing for maximumeffectiveness.This placementconsidered the

possibilityof a tip stallcondition,wherebythewing tip would stall beforetheroot. By

placingthe aileronsapproximatelytwo-thirdsof the way out alongthe wing, neithertip

stall norroot stall wasfavored. Thus,if eitherthe root or the tip doesstall, therewill still bea substantial portion of the aileronin thefreestream to sustainits effectiveness.Thus,

theinboardlocationof the aileronwaschosen as 1.75feetandthe outboardlocationwas

chosen as3 feetto give a roll rate of 20.5degrees/second.

Oneweakness of the aileronswasthatat low speeds andon airplaneswith large

spans, induceddragwasdominant. The aileronstendto produce"adverseyaw", or yaw

dueto the down-goingaileronproducingmoredragthanthe up-goingaileron. This

tendency of theaileronstopull theaircraftawayfrom the turn canbecounteredwith

simultaneous applicationof a rudder.

Table 7.5.1: Lateral Control Parameters 1.25 ft yl (inboard distance) 1.75 ft ),2 (outboard distance) Roll rate (p) 20.5 deg/sec 1.25 ft span chord 1.5 inches 0.155 Clda 7.6 Directional Control: The directional control of the aircraft was created by deflecting the rudder on the vertical tail. The size of the rudder was driven by its need to provide a sufficient yawing moment to overcome the adverse yaw created by the ailerons. The control power for the rudder was investigated in Figure 7.6.1 for different rudder sizes and vertical tail sizes.

Figure 7.6.1: Control Surface Area Ratio vs. Control Power 0.1 - I 0.i -. m .. --, -0.1 I " -0.2 ---I-- Sv/S = 0.015 "'" -0.3 -: m _ ---0-- Sv/S = 0.045 o -0.4 (3.. Sv/S = 0.078

-0.5

I • -e-- Sv/S = 0.11 t- O -0.6 I I -O.7 I I -0.8 0.1 0.2 0.3 0.4 0.5 0.6 0 0.7 Rudder Area / Vertical Tail Area The yawing moment needed to counteract the adverse yaw due to the ailerons was calculated. The necessary counteracting yaw moment of the rudder requires a rudder control power of at least -0.08 rad -1 at a maximum deflection angle of + 15 degrees. The top horizontal dashed line in Figure 7.6.1 shows that no parameters above the line can be considered. This control power results in a minimum ratio of rudder area to vertical tail area of 0.42. This fact can be seen by the vertical dashed line in Figure 7.6.1. Ratios to the left of this line are not valid since the rudder sizes will not provide sufficient yaw moment. When the minimum ratio was chosen, a rudder area of 0.12 ft 2 resulted. The previously chosen vertical tail area of 0.3 ft 2 (S v/S = 0.46) resulted in a sufficient control power for overcoming the adverse yaw, providing a Cndr value of -0.14 rad -1. This value was significandy lower than previous Aeroworld airplanes, however this may be due to the fact that many of them did not have ailerons. The rudder for those airplanes had to have more control power in order to turn the airplane, coupled with the wing dihedral.

The control power obtained for this particular aircraft should be quite sufficient.

The minimal vertical tail area was not chosen because of the uncertainties involved. This results in a large vertical tail that does not look like it belongs to this particular airplane design, making the aircraft appear "awkward". This decision goes against the design goal of an aesthetically pleasing airplane, one of the objectives which helped determine the stabilator size. However, the greater uncertainties involved in directional stability and control requires a conservative tail size in order to ensure stability. This decisionseems justified.

Clearly,evena good-lookingairplanewasnot

worth muchif it cannotbecontrolled.

Table 7.6.1: Directional Control Characteristics Sr / Sv 0.046 Sr 0.30 ft 2 -0.103 Cndr dr max + 15 de_rees 7.7 Control Mechanisms: The Elite utilizes three different control surfaces: a rudder, an all-moveable tail, and ailerons. The rudder will have a maximum deflection of + 15 degrees and the all- moveable tail will be able to deflect up to + 12 degrees. The tail will be mounted at an incidence angle of -2 degrees for zero deflection at cruise condition.

Control will be provided by way of control rods connected to each surface and a series of servos. Plastic control rods will be used due to their simple operation and flexibility. The plastic rods maneuver freely within the nylon tubing, thus allowing smooth movement of the control surfaces. The flexible rods will enable them to be bent around other components such as the batteries.

Each control rod will be connected to the control surface by way of a control horn.

The control horns have adjustable connection joints so that the surfaces deflections can be altered. One servo will link the control of both the rudder and the nose gear. The ailerons and horizontal tail will both operate on separate servos. All control rods will be internal to the aircraft in order to help decrease any unnecessary drag.

7.8 References: 7.1 Nelson, Robert C., Flight Stability and Control, New York, MCGraw-Hill Book Company, 1989.

7.2 Simons, Martin, Model Airplane Aerodynamics, Great Britain, Argus Books Limited, 1978.

7.3 MCCormick, Barnes W., Aerodynamics. Aeronautics. and Flighl Mechanics, New York, John Wiley & Sons, 1979.

7.4 Avis, Daniel, Folaxan computer code for take-off rotation, February, 1994.

8 AIRCRAFT PERFORMANCE

8.1 Requirements and Objectives Requirements: 1. capable of a sustained, level 60 foot radius turn at a speed of less than 30 feet/second 2. rough field characteristics a. adequate taxi and runway handling characteristics b. able to climb to a height of 50 feet within 200 feet of brake release c. maximum take-off distance of 60 feet 3. able to fly to nearest alternate airport and loiter for one minute Objectives: 1. minimum cruise speed of 60 feet/second 2. maximum velocity greater than 80 feet/second 3. sufficient range to service aU Aeroworld airports 4. endurance consistent with target range, cruise and loiter speeds 5. maximum take-off distances a. rough field, 42 feet b. improved runway, 28 feet 6. handling qualities consistent with private/sport recreational aircraft 8.2 Summary of Performance The upper-class market at which The Elite is aimed demands an aircraft that will not only be aesthetically pleasing, but will exhibit a high level of performance for its class of aircraft. The performance requirements outlined in the Design Requirements and Objectives drove the design of The Elite. Specifically, the required take-off distance of 28 feet, the cruise velocity of 60 feet/second and a maximum velocity of 80 feet/second were primary drivers in the choice of propulsion system. The desire to service all airports in Aeroworld was also a major concern as was the ability to execute a 60 foot radius turn at 28 feet/second. Table 8.1.1 illustrates the performance specifications of The Elite.

7O Table 8.2.1: Performance Specifications for The Elite Aircraft Speed Performance: 60 ft/s 24.8 ft/s Minimum Velocity (Stall Velocity) at WMTO _rance: 32900 fl (at V=36.9 ft/s) Maximum Range at WMTO 14.9 minutes Endurance at Maximum Range at WMTO 30500 feet (at V--49.6 ft/s) Design Range at WMTO 10.3 minutes Endurance at Design Range (WMTO) 27268 ft (at V=60 ft/s) Cruise Range at WMTO Climb and Glide Performance: Maximum Rate of Climb at WMTO 16 ft/s (at V=36.5 ft/s) Maximum Climb Angle 29.9 degrees (at V=30 ft/s) Take-Off Performance: Take-Off Distance at WMTO 25.54 ft 8.3 Take-Off Performance In order to service all of the runways in Aeroworld, the DR&O set a maximum take-off roll on an improved runway at 28 feet. This requirement influenced the choice of the Astro 15 motor combined with a Zingali 10-8 three-blade propeller as the propulsion system. The FORTRAN program TAKEOFF (Ref. 8.1) was used to determine the distance required by The Elite to lift-off. The TAKEOFF program, which uses a numerical integration routine to compute take-off roll, indicated that the distance required by the aircraft to take-off was 25.5 feet. This value was determined using pessimistic values of a Cl-max of 1.00, a weight of 4.88 pounds, and a rolling friction coefficient of 0.19. In order to achieve these values, TAKEOFF assumes the aircraft runs its propeller up to maximum rpm and then releases brakes. Since the RPV is not equipped with brakes it is anticipated that the take-off roll will be slightly longer than predicted by the program.

Figure 8.3.1 illustrates the results of an investigation of the dependence of take- off distance upon manufacturing imperfections. Because of the uncertainty in the maximum coefficient of lift of a manufactured airfoil, the variation of take-off distance with weight and maximum coefficient of lift were investigated. Figure 8.3.1 indicates that a maximum take-off weight of approximately 5.2 pounds is the most the technology demonstrator may weigh to satisfy the take-off distance requirement. This result comes from the belief that it is possible to achieve a maximum coefficient of lift of 0.95 for the airfoil as opposed to the design value of 1.14.

Figure 8.3.1: Effect of Manufacturing Imperfections upon Take-Off Performance

J

0.)

30- e" Cl-max=.98 28 't take-¢,ff limit °m Cl-max=.95 Cl-max=.9 _._ ,._nt loca Lion 2O _ _'" of dq _ign 4.2 4.4 4.6 4.8 5 5.2 5.4 5.6 Aircraft Maximum Gross Take-Off Weight (lbs) 8.4 Range and Endurance The Design Requirements and Objectives Document specified a range sufficient to service all airports in Aeroworld allowing for diversion to an alternate airfield including a one minute loiter. The range specified in the DR&O was found by determining the distance to each airport and its closest alternate and adding 1 minute of loiter time at a velocity of 30 ft/s. Using this method, a design range of 30500 feet was deemed necessary to serve all airports in Aeroworld. This study used a computer program (Ref 8.2) written to determine range and endurance as a function of velocity.

Upon completion of the study it was determined that it was not possible to achieve the range specified in the DR&O at the cruise speed of 60 ft/s with the current propulsion system. A study was then undertaken to determine how much of Aeroworld was serviceable with a range of 27268 ft at the cruise velocity of 60 ft/s. After examining every possible route (including distances to alternate airstrips and loiter i_ne) in Aeroworld, it was determined that only six routes were not serviceable at a velocity of 60 ft/s. At the design cruise speed of 60 ft/s, The Elite can service 94.3% of all possible routes in Aeroworld. The six routes that cannot be serviced at a speed of 60 ft/s can be serviced at a minimum speed of 49.6 ft/s. This compromise was deemed adequate and it was decided not to reduce cruise speed or reconsider the propulsion system.

Figure 8.4.1 illustrates the relationship between range, endurance and velocity at the maximum take-off weight of this aircraft. This figure indicates that the maximum range for The Elite is 32900 ft at a speed of 36.9 ft/s. Also indicated on this plot is the location of the maximum endurance and the design range and the cruise range for this aircraft.

Figure 8.4.1: Variation of Range and Endurance with Velocity 34000 21 33000- 32000- 31000- 30000- 29000- e_ 28000- 27000-

\

26000- 25O00-.

24000_ The effect of weight on range was also investigated in Figure 8.4.2. This figure shows the linear dependence of range on weight and the minimal variation of the range for this small general aviation aircraft. This minimal variation is due to the small weight of the payload carried by The Elite.

Figure 8.4.2: Dependence of Range and Endurance Upon Payload 27490- -7.65 ra 0 Passengers

-,.,, .7.645

27470 0 3 Passengers .7.64 : 6 Passengers 27450 .7.635

\

.7.63 E 27430 ID .7.625 27410 .7.62 .7.615 27390 • 0 Passengers .7.61 • 3 Passengers 27370 .7.605 A 6 Passengers 27350 .7.6 .... I .... I 4.82 4.83 4.84 4.85 4.86 4.87 4.88 4.89 Payload (lbs) 8.5 Power Required and Power Available Figure 8.5.1 shows the power required and power available curves at varying motor voltage settings obtained using Ref. 8.3. The maximum velocity of 71.7 ft/s is evident at the far right intersection of the power required and power available curves.

This velocity is the maximum velocity at which the aircraft may fly and still maintain steady, level flight. Also indicated on the plot is the voltage setting of 12.9 volts necessary to maintain steady cruise and the voltage settings for maximum range (8.0 volts) for the maximum velocity (15.6 volts) and for the velocity necessary to service all Aeroworld routes (10.5 volts). In addition to the higher maximum velocities induced by higher voltage settings, the climbing ability of the aircraft also improves because of the dependence of rate of climb on the difference between power available and power required at specific velocities.

Figure 8.5.1: Power Required and Power Available Curves 00000000000000 O0 [] [] [] e_ o o

, i 1 !

.... I .... I .... I .... I .... I ....

20 30 40 50 60 70 80 Velocity (if/s)

8.6 Climb and Gliding Performance

Using the power available and power required curves acquired in the previous section the rate of climb for the aircraft was determined by simply taking the difference between the power curves and then dividing by the weight of the aircraft (Ref 8.4). Using this method, the maximum rate of climb for The Elite was determined to be 16 ft/s at a velocity of 36.5 ft/s. This velocity is slightly greater than the take-off velocity meaning that the aircraft will be flying at or close to the maximum rate of climb in the portion of the flight regime where good climbing ability is necessary. Assuming a cruising altitude of 25 feet and a climb-out angle of 23.67 degrees at the maximum rate of climb The Elite will reach the cruising altitude of 25 feet in 1.6 seconds while covering a ground distance of 57 feet. Including the ground roll distance, this allows the aircraft to be in cruise approximately 67 feet before encountering the first turn in the Loftus Center. This rate of climb also ensured that the aircraft would be able to satisfy the design requirement of being able to climb to an altitude of 50 feet within 200 feet of brake release. Assuming a take-off roll of the design objective length of 42 feet and a maximum rate of climb, the aircraft would be able to climb to a height of 50 feet within 156 feet of brake release satisfying this design requirement.

Examining the glide performance of The Elite simply involved knowledge of the maximum lift-to-drag ratio of the aircraft (Ref 8.5). By inverting the maximum lift-to- drag ratio and taking the arctangent of the result, the minimum glide angle was obtained.

This method resulted in a minimum glide angle of 3.87 degrees. This value is important in an engine out scenario. If the single engine of The Elite were to fail, a gentle glide slope was desirable to ensure a safe landing. If, perhaps, the engine were to shut down at an altitude of 25 feet, at the minimum glide-slope the RPV will cover 370 feet at the minimum glide angle before touching down.

8.7 Turn Performance The Elite is required to perform a 60 foot radius turn at a speed of less than 30 ft/s as specified in the DR&O. In order to satisfy that requirement, the aircraft will have to bank an angle of 23.9 degrees and make the turn at a speed of 28 ft/s. Using the formulas supplied by Ref 8.6 the radius determined for the turn was 55 ft and the g-factor was found to be 1.1g. A maximum bank angle of 70 degrees was found using a maximum load factor of 3 as provided by the structural design of The Elite. This maximum bank angle results in a turn radius of 41 feet at the cruise speed of 60 ft/s.

8.8 References

8.1 Batill, Dr. Stephen. TAKEOFF FORTRAN program. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

8.2 Le, Tuan. A FORTRAN program to compute range and endurance, 1994.

8.3 Batill, Dr. Stephen. PAVAIL FORTRAN program. Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1994.

8.4 McCormick, Barnes W. Aerodynamics, Aeronautics and Flight Mechanic_. New York: John Wiley and Sons Inc., 1979. p. 432.

8.5 Jumper, Dr. Eric J. "AE441 - Flight Mechanics/Intro to Design Class Notes - Aircraft Performance." Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1993.

8.6 Jumper, Dr. Eric J. "AE441 - Flight Mechanics/Intro to Design Class Notes - Turn Performance." Department of Aerospace and Mechanical Engineering, The University of Notre Dame, 1993.

STRUCTURAL DETAIL DESIGN

9.1 Requirements and Objectives Design Requirements: 1. Fuselage must contain 6 passengers and avionics.

2. Structure must be designed so as to ensure the survivability of passengers and radio components in a crash from any flight condition.

3. Structure must allow for easy access to avionics and propulsion system so that complete system installation can be accomplished in no more than 20 minutes.

4. All propulsion system batteries must be placed in the main wing carry- through structure.

Design Objectives: 1. A lightweight structure of less than 4.88 lbs so as achieve performance objectives.

2. An aerodynamically efficient aircraft with aesthetic appeal so as to be marketable in the higher end of the general aviation market.

3. A structure capable of withstanding flight load factors of no greater than 3.0 and no less than -2.0 and ground loadings of up to 3.0 G's with a factor of safety no less than 1.5.

4. Readily accessible avionics and propulsion system components allowing installation in less than 20 minutes.

9.2 Main Wing Spar Design To determine the best structural design for the wing, an extensive trade study was performed on six different balsa spar configurations for the cross section of the wing.

The loads used in this test subjected the wing to three times the landing load distribution and at the same time put three times the weight of the aircraft on the main gear. This should provide a worst case scenario that will not be experienced in the operation of the aircraft. The shear and bending moment graphs associated with this loading configuration are shown in Figure 9.2.1.

The cross section was idealized into four lumped normal stress carrying booms and four shear panels (see Figure 9.2.2). Table 9.2.1 presents a summary of the analyzed configurations. One thing to note is that the fifth and sixth configurations use graphite tape on the bottom main spar. The modulus of elasticity of the tape enforced spar was 11,000,000 lb/in 2 versus the un-enforced spar 800,000 lb/in 2. In effect, this stiffens the total beam against a tip up deflection.

Figure 9.2.1: Limit Loading of Wing a) shear b) moment 16- 14J .

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.

0 0.5 1 1.5 2 2.5 3 3.5 4 0 0.5 1 1.5 2 2.5 3 3.5 4 Spanwise Location, (ft) Spanwise Location, (ft) a) b) Figures 9.2.1 a) and b) show the shear and bending moment diagrams for the wing in the limit load configuration previously described. The jump in the shear diagram is due to the landing gear loads placed 4.5 inches from the root.

Figure 9.2.2: Idealized Cross-Section #1 #4 #2 #3

(o,o)

Table 9.2.1: Idealized Cross-Section Coordinates and Dimensions z-loc y-lot Different Tested Configurations--Spar Areas (in 2) 1 2

lump # (in) (in)

3 4 *5 *6 0.6875 0.03125 0.0469 0.0625 0.0625 0.03125 0.0469 0.0625 0.03125 0.03125 0.03125 0.03125 0.03125 0.03125 0.0 -0.3125 0.03125 0.0469 0.03125 0.0625 0.0391 0.0391 2.547 0.0 0.0352 0.0352 0.0352 0.0352 0.0352 0.0352 The trade study was performed with a code that Jonathan Fay wrote for AE346- Aircraft Structures. The code was validated in that class. The trade study attempted to find the configuration with the lowest weight, highest stiffness, and lowest stress. The weight of each design was calculated from the total volume of the main wing spars multiplied by an average density of balsa. The stiffness of the beam was evaluated on the basis of how far the tip would deflect under a constant load distribution.

wE'

5'iP = BE'---I" (ref. 2) Finally, the stress in each lump was outputted from the computer code under the loading conditions described above. These three factors were combined into a single figure of merit, Z, based on their relative importance to the design. The weight of the design received a weighting of 3 while the tip deflection received a weighting of 1.5 and the stress in the #1 lump received a weighting of 1. The result of this study is shown graphically in Figure 9.2.3.

Figure 9.2.3: Comparison of Tested Cross-Sections 0.25.

0.2- CD 0.15 O _ 0.1 u_ 0.05 g i I ! i | I l | 1 i i i i i i i m m ! m 2 3 4 5 6 Configuration * These two analyzed cross-sections contain a strip of graphite re-enforcing tape along the length of the third lump (i.e. the bottom main spar).

where the Figure of Merit, Z, Z 3 × Weight + 1.5 x Tip Deflection + Stress in #1 lump / (-4000) Since it is desirable to have the lowest possible weight, with the smallest tip deflection, and lowest stress, the best design is represented by the maximum value of the figure of merit. Thus the sixth configuration was chosen as the design for the wing.

9.3 Main Wing Rib Spacing A trade study was conducted to determine the proper rib spacing for the wing. To save on structural weight, it is desirable to space the ribs as far apart as possible. Two main factors played a role in determining the final value of 4" for the rib spacing. The first major factor in determining the rib spacing was the shaping of the monokote. There is a point when the monokote sags between the ribs enough to hurt the wing performance aerodynamically. By examining past years wings, 4" was found to be the largest value at which a reasonably consistent airfoil shape can be maintained.

The second factor was the buckling of the wing spars. The ribs must be close enough together to prevent the main wing spars from buckling under the worst case scenario loads. The buckling characteristic of the spars was determined using the a pinned end approximation in conjunction with the stress analysis output by the computer code. From reference 2, the buckling length of a rod pinned at both ends is given by: Lbuckle "- n_ (ref 2) P is the applied force to the end of the rod. To determine the value of P, the normal stress in each spar given by the computer code was multiplied by the spar's cross- sectional area. To prevent buckling under the worst case scenario loading conditions, the ribs would have to be spaced 1.5" apart at the sections near the wing root. This was an unacceptable solution because it added too much weight to the aircraft. Ideally, a spacing of 4" would be used since that is the largest value possible due to aerodynamic (wing shaping) concerns. To solve the buckling problem, some balsa webbing was added to the inboard four sections of the wing. The webbing effectively increases the mode of buckling (n in the above equation) that the spars would undergo, thus increasing the length at which buckling would occur. The final webbing configuration to prevent buckling at the limit loads with a 4" rib spacing is shown below. The inboard most section has a full sheet of 1/16" balsa running across the main spars. The next three sections of the wing use 3/16" square balsa pegs as the webbing.

Figure 9.3.1: Webbing Between Main Spars (Wing Front View) root _ _op l rib#1 rib #2 rib #3 rib#4 / spar no webbi_ beyond thisrib 9.4 Fuselage Structural Design The primary design objectives for the fuselage design were low weight, low drag, aesthetics, strength, and system access. After consideration of the initial concepts, a rounded fuselage composed of longerons and bulkheads was chosen because it satisfied each of the above.

Of primary concern were the sizes of the longerons and the bulkheads. Two studies were conducted, the first of which determined the dimensions of the longcrons needed to provide the necessary strength while minimizing the weight. The software developed to conduct the wing structure analysis was modified to model the fuselage as a right circular cylinder with a 4" diameter (the average diameter of the tapered fuselage) formed by four longerons (see Figure 9.4.1). Though spruce and basswood were considered initially, balsa proved to be more than adequate to withstand the loads while minimizing the weight, a primary design concern. Figure 9.4.2 depicts the idealized cross section composed of four lumped booms and shear panels. Table 9.4.1 lists the three configurations examined under flight loads during cruise and ground loads during landing (see Figures 9.4.3 and 9.4.4).

Figure 9.4.1: Fuselage Model for Stress Analysis Figure 9.4.2: Idealized Cross-Section LUMP 1 PANEL 0 z 1 N.A Table 9.4.1: Idealized Cross-Section Coordinates and Dimensions z-loc Configuration--Longeron Dimensions (in)

lump # (in) 2 3

2.0 0.0 1/4 x 1/8 1/4 x 3/16 1/4 x 1/4 0.0 2.0 1/4 x 1/8 1/4 x 3/16 1/4 x 1/4 1/4 x 1/8 1/4 x 3/16 1/4 x 1/4 1/4 x 1/8 1/4 x 3/16 1/4 x 1/4 Figure 9.4.3: Fuselage Ground Loading a) shear b) bending moment t- 2.5 --.

_._ 2- 1.5-

W_ o._

o 1-

_ -4

i ¢'

o.5J___:

g_ -6

-0.5

g -a

u_ -1.5 -10 0 5 10 15 20 25 30 35 40 u_ 0 5 10 15 20 25 30 35 40 X-Location, (inches from nose) X-Location, (inches from nose) a) b) Figure 9.4.4: Fuselage Loading at Cruise a) shear b) bending moment ,-,3- 5.-.

z o--: f_:

o

..... ° 10 m ....

oil

CO -15 y

-1 .i _

=m -20

-_ -2 -!

-25- [33 P!

-30 U-_ 3 0 5 10 15 20 25 30 35 40 0 5 10 15 20 25 30 35 40 X-Location, (inches from nose) X-Location, (inches from nose) a) b) The shear diagrams were obtained by discretely integrating the weight of the aircraft from the nose to the tail and including the point loads due to the gear and the lifting surfaces. Similarly, the moment diagrams were calculated by discretely integrating the shear along the length of the aircraft. One thing to note is that the moment charts were extremely sensitive to the placement of the major forces. Even a relatively small (0.5 in) movement in the landing gear or lift forces would prevent the moment diagrams from returning to zero at the tail. Once these loading distributions were known,

theresultingstresses in thelongerons couldbecalculatedso asto selecttheminimum

longeronsizewhich would sustaintheloadswhile satisfyingthe design objective of a

margin of safety no less than 0.5. Table 9.4.2 summarizes the results of the stress analysis (Ref. 9.1 provided t_fail for balsa).

Table 9.4.2: Longeron Stress Analysis Results Configuration margin of safety fffail (psi) ffm_x (psi) 5337 2914 0.8 2 5337 1942 1.7 5337 1457 2.7 The study indicated configuration one (1/4" x 1/8" longerons) would achieve the design objectives of minimizing the weight and attaining a margin of safety of at least 0.5. There were two additional concerns, though, which were addressed before arriving at a final longeron design. First, the fuselage tapers considerably fore and aft of the wing, a feature of the design unaccounted for in the stress analysis. As a result of the taper, the moment of inertia of the longerons along the top and bottom of the fuselage decreases as the longerons approach the neutral axis. Recalling that stress is inversely proportional to moment of inertia, and also considering that the top and bottom longerons bear the greatest stresses in the fuselage structure, it was apparent that configuration one may not be sufficient to withstand the loads with a margin of safety of at least 0.5. Second, a calculation of the buckling length of the longerons when _ = Crnax indicated the bulkheads would have to be spaced less than 1 7/8" apart to prevent buckling. This was undesirable because the need for so many bulkheads largely negated any weight savings derived by selecting the longerons with the least cross-sectional area. Clearly, a means of strengthening the top and bottom longerons with a minimum weight penalty was needed.

Carbon fiber tape proved to be the solution. The addition of strips of carbon fiber tape epoxied along the top and bottom longerons greatly strengthened the longerons.

With a modulus of elasticity of 11.0E6 lb/in 2, the combination of the longeron and 1/4" x 7/1000" tape can withstand any tensile load the fuselage will experience (the modulus of elasticity of the composite longeron was determined in Section 9.2). Thus, the final longeron design was 1/4" x 1/8" balsa stock with carbon fiber tape epoxied along the top and bottom longerons. Four additional longerons were added to help support the Monokote coveting. Since the four primary longerons bear the loads, lightweight 1/16" x 1/16" balsa stock was chosen for the secondary longerons. The addition of the carbon fiber tape increased the buckling length to 5", though the bulkhead spacing aft of the

wing wassetat 2 3/4" andthe bulkheads fore of the wing werespaced 2 3/16" apartso as

to supportthe secondary longerons andpreventtheMonokotefrom sagging. Within the

wing carry-throughstructure,the bulkheads werespaced to accommodate the wing

attachment structure(seeFigure9.4.5).

Figure 9.4.5: FuselageStructure

The second studyinvolved sizingof thefuselagebulkheads.In particular,an

analysisof the shear flow throughthe bulkheads wasconductedso asto determinethe

minimumcross-sectional thickness necessary to withstandthe loads. Minimizing the

thickness will in turn minimize theweight soasto attainthedesignobjectiveof a

lightweightstructure.

The bulkheadstudyemployedthefuselagemodelandidealizedcross-section used

in thelongeronanalysis(seeFigures9.4.1and9.4.2). Forthis study,however,the shear

flow in eachpanelwascalculatedgiventhe flight andgroundloads(recallFigures9.4.3

and9.4.4). With the bulkhead modeledasa sliceof a thin-walledmember,the average

shearcouldthenbe estimated (seeFigure9.4.6andRef. 9.2).

Figure 9.4.6: Bulkhead Model for Shear Analysis V

/¢Zd

// ////-

Note that the thickness is measured inside of the notches for the longerons and not from the outer circumference. Doing so increased the overall thickness so as to provide for stress concentrations at the cuts. As with the longerons, balsa was chosen due to its high strength-to-weight ratio. The three configurations examined were t = 1/8", 3/16", and 1/4". As with the longeron study, low weight and a margin of safety of at least 0.5 were the primary figures of merit. Table 9.4.3 summarizes the results (Ref. 9.3 provided tfail for balsa)..

Table 9.4.3: Bulkhead Shear Analysis Results Configuration tfail (psi) tm_ (psi) margin ofsa_ty 1 300 24.32 11.3 2 300 16.21 17.5 3 300 12.16 23.7 The above results indicate that the first configuration (t = 1/8") would attain the design objectives. However, due to concerns about the stress concentrations at the notches, configuration two (t = 3/16") was selected. A further concern was the orientation of the grain of the balsa; the bulkhead would be weak at the points where the grain was oriented radially (see Figure 9.4.7). Therefore, the material was changed from sheet balsa to 3-ply balsa sheets. The grains within the plywood balsa are mutually perpendicular, thus eliminating the weaknesses due to unidirectional grains.

Figure 9.4.7: Balsa Sheet versus 3-Ply Balsa a S t s s \ The f'mal consideration was the thickness of the balsa stock from which to fabricate the bulkheads. In order to minimize weight without sacrificing strength, several thicknesses were used for the bulkheads depending upon the estimated load on each. For instance, 1/8" stock was chosen for the firewall due to the sizable loads placed upon it by the power plant and nose gear. Likewise, the bulkheads within the wing carry-through structure are sturdy 3/32" 3-ply balsa as are the bulkheads fore and aft of the stabilator hinge. The remaining bulkheads are 1/16" thick (see Figure 9.4.5).

9.5 Wing Carry-Through and Fuselage Mating Design The wing carry-through design warranted particular attention due to the many design objectives it needed to satisfy. First and foremost, the structure had to be lightweight yet strong enough to withstand the flight load extremes as well as the ground loads experienced by the main gear. Furthermore, the placement of the avionics and propulsion system batteries within the carry-through structure posed its own difficulties, for in order to attain the design objective of complete system installation within 20 minutes, the wing would have to be simple to remove and mount. A bolt system was chosen due to the ease of removal and remounting that it offered (see Figure 9.5.1).

Though such a design demands careful cunstruction for proper mating of the wing and fuselage, it offers distinct advantages over a rubber band system or a "plug-in" system by which the sing spars detach from the carry-through structure. Bolts offer an internal mounting system without the drag of external rubber bands and wooden dowels, while it also avoids the weight penalty incurred in strengthening the wing spar joint in a plug-in system.

The wing is connected to the fuselage by four nylon bolts. The bolts are screwed through spruce blocks epoxied to the wing's main and secondary spars and continue into

spruceblockssandwiched betweenfuselagebulkheads.The wing sitsin a saddleformed

in the bottomof the fuselageby thebulkheads andthin balsasheeting(seeFigure9.4.5).

The bottomlongeronterminates at thewing's leadingedgeandresumesatthe trailing

edgesothatthe wing canbeplacedflush with the fuselage.As a result,theload paths

run form the wing, throughthe carry-through structurebulkheads, andtheninto 3/16" x

3/16" sprucebeams runningalongtheinsideof the bulkheads.The two beamsservethe

dualpurposeof providing strength while providing a platformto mountthe avionics.

Thewing sparsarejoined atthefuselagecenterline andwrappedwith a thin strip

of fiberglasscloth andepoxy. The bottomof the wing box is sheeted soasto providea

floor to which thepropulsionsystembatteries arevelcroed. The top of the wing box is

partially sheeted wherethe wing contacts thebottomof the sidesof the fuselage.Partial

top sheetingin the aft sectionof thewing box providesa mountfor the aileronservo.

Figure 9.5.1: Wing Mounting System \ \ \ \ \ / \ \ 1/4" x I/4" / \ spruce / \ / \ \ \ \ \ \ 1/2" x 1/2" basswood wing spars nylon bolts 9.6 Empennage Design The tail structure was driven by the design objectives of a lightweight structure with the strength to withstand the flight loads. Though ground loads are not of particular concern in designing the empennage, the flight loads are compounded by the additional loads induced by control surface deflections. A further concern is the sensitivity of the centerof gravity locationto additionalweight at thetail.

With theseconsiderations in mind, the empennage wasdesigned with a simple

structure.Theverticaltail is merelya flat platewith a trussstructurewhereasthe rudder

is 1/4"sheetbalsawith holesdrilled to reducethe weight(seeFigure9.6.1). Two nylon

hingesandaninternalwire actuatorconnecttherudderto the verticaltail. A 1/4" x 1/4"

balsabeamextendingto the bottomlongeronserves asthe mainsparof the verticaltail,

whereas1/8"x 1/4" balsaformstheleadingedgeandhorizontalmass members.The

diagonaltrussmembers are1/16"x 1/4" balsa. Empiricaldatafrom previousdesigns

provideda basefor decisions concerning the dimensions of thevariousmembers.

Thedecisiontoincorporatea stabilatorratherthana horizontaltail/elevator

combinationwasdrivenprimarily by performance considerations, thoughthe all-

moveabletail offeredstructuraladvantages aswell. Thoughmountingthe stabilator

requireda morecomplexdesign,weightwassavedby eliminatingthe nylon hingesand

additionaltrailing edgestructure needed for elevators.The designobjectiveto createan

aesthetically pleasingaircraftwasthebasisfor sweeping the stabilator15° aft.

The primarycomponent of the hingedesignis a lightweight 1/4"diameter12"

long carbonfiber rod which canwithstandthe stabilator'storsionalandtransverse loads

(seeFigure9.6.2). Soasto minimizethe hingemoment,the centerlineof therod was

/

/

/ fuselage longeron Figure 9.6.1: Vertical Tail Structure designed to pass through the intersection of the quarter chord and the mean aerodynamic chord. Recall that for a symmetric airfoil, the coefficient of moment about the quarter

chordis zero. Hence,thehingemomentwill benearzero(the additionof a smallfillet

betweenthe inboardrib andthefuselagewill producea smallmoment). Therod serves

asa partial sparfor the stabilator, extendingto thethird rib. Holesdrilledin thethree

inboardribs serveastheattachment points. A roundedpieceof 3/16" x 3/16" balsa

servesasthe leadingedge,whereas 3/8" x 3/32" balsaformsthe trailing edge.Two

forward 1/4"x 3/16" sparsandtwo aft 1/8" x 1/8" sparsserveto strengthen the tip of the

stabilatoroutboardof therod andsupportthe Monokotecovering. The stabilatormounts

to the fuselagevia two 1/2"x 1/2" spruce blocksheldin placebetweenthe two aft

bulkheads (seeFigure9.6.3).

Figure 9.6.2: Stabilator Structure i/2" x 1/2" hardwood

I

..... _Jl ...... _...............

_-12" x 1/4" \ diameter carbon I 1/4 _ mac = 6" Figure 9.6.3: Stabilator Mounting System i/2" x 1/2 spruce control horn carbon fiber rod 9.7 Aircraft Loading Based on the ultimate loads placed on the fuselage and wing when they were structurally designed, a good estimate for the limit load factor for positive angles of attack is 3, while -2 should roughly approximate the allowable load factor for negative angles of attack. These values provide the pilot with a reasonable operation's envelope while still adequately safeguarding the smactural integrity of the aircraft. Figures 9.7.1 and 9.7.2 show the velocity/load factor envelope and the corresponding velocity/angle of attack envelope for our aircraft respectively.

Figure 9.7.1: V-n diagram for Elite Aircraft .

Limit Load b CLmax +3 deg "_ ' t- O -1 -3.5 deg "h, -2 "_ Limit Load " -3 ' ''' I' ''' I'' '' I'' '' I ''' '1'_ 8O 0 10 20 30 40 50 60 70 / Velocity, (ft/sec) Vma_ Notice in the V-n diagram that the extreme allowable angles of attack at Vmax are a modest 3 and -3.5 degrees. This could present some difficulties in preventing the pilot from over stressing the aircraft at the higher flight speeds. Figure 9.7.2 shows the full dependence of allowable angle of attack on velocity. Once the airplane gets above 36 ft/sec the pilot is no longer allowed to stall the aircraft without the danger of damaging the aircraft. In fact, if the n=3 and n=-2 load factor curves are carried out to 150 ft/sec, as could be encountered in a dive, the allowable angle of attack range narrows to -0.18 degrees to - 1.87 degrees.

Figure 9.7.2: Angle of Attack and Velocity Envelope for Elite Aircraft stall 10- CD CD "O v O c_ O -5 c-- -10 Vmax stall -15 10 20 30 40 50 60 70 80 90 100 Velocity, (ft/sec) 9.8 Landing Gear Design The landing gear system employed is the standard RPV steel strut gear in a tricycle pattern. The 1/8" steel struts, main gear basswood blocks, and nose gear mounting brackets will withstand the ground loads during takeoff and landing from unimproved fields. However, there are other concerns surrounding the landing gear that should warranted some attention.

First, the gear must be long enough to provide adequate ground clearance for the propeller. Second, in a tricycle formation, the main gear must be behind the center of gravity, but not so far behind as to prevent the aircraft from rotating at takeoff. The rear gear should also have a large enough spacing between them to prevent the aircraft from tipping over during ground maneuvers. Lastly, the landing gear must attach to a very sturdy part of the aircraft to prevent structural damage if the aircraft hits the ground during a landing attempt. To meet this last requirement, secondary wing spars placed 5" behind the leading edge support 1" thick basswood blocks to which the main gears mount (see Figure 9.8.1). These secondary spars are made of 3/16" square spruce and allow a stable attachment point for the gear. The use of spruce should also prevent the gear from tearing out of the wing.

Figure 9.8.1: Main Gear Mounting System wlng spars i 1" x i" basswood 5/32" wire f / \ / \ \ \

t i

2" diameter wheel I " / \ / \ / \ / \ j/ Table 9.8.1 below summarizes the landing gear properties that meet all the above stipulations for our aircraft.

Table 9.8.1: Summary of Landing Gear Properties Material Diameter Nose Gear Main Gear Main Gear Gear Length Position Position Spacin_ steel 1/8" 4.125 in 15.5 in 21 in 6 in In terms of the overall design, the landing gear did not receive much attention and is one area that could probably use a more detailed analysis.

9.9 References

9.1 Solbreken, Gary L. "'Balsa Wood Breaking Experiment" (memo to Dr. S.

M. Batill). 30 July 1992.

9.2 Beer, Ferdinand P. & Russell Johnston, Jr. Mechanics of Materials. New York: McGraw-Hill, Inc. 1992.

9.3 Wood Engineering Handbook. ed. Forest Products Laboratory.

Englewood Cliffs, N.J.: Prentice Hall. 1990.

10 ECONOMIC ANALYSIS

10.1 Requirements and Objectives Design Objective: 1. Make The Elite affordable to the upper-middle class general aviation market.

10.2 Cost Breakdown The total cost per aircraft was a function of the f'txed subsystem costs, the raw materials costs, and the manufacturing costs. The complete cost breakdown is found in Table 10.2.1. The total cost of the fixed subsystems was $462.58. The cost of the raw materials was estimated as $160. Manufacturing costs were estimated at $2190.

Included in the manufacturing costs were the personnel costs, the tooling costs, and the waste disposal costs. The personnel costs are based on 180 hours of manufacturing time, a conservative estimate based on the complexity of our design versus that of previous designs. Waste disposal was approximated at 1.5 lbs, also a conservative estimate based on the prediction of excess material after cutting out the circular fuselage and swept empennage. The total cost of the subsystems, raw materials, and manufacturing was $2812.58. Assuming an overhead factor of 1.4 and a 12% profit, the total cost of The Elite was $4410.13.

A breakdown of the three main components affecting the total cost is found in Figure 10.2.1. Manufacturing costs make up the largest percentage of the total cost of the aircraft (78%). The subsystems make up 16% and the raw materials only 6%. Thus, to decrease the total cost of the aircraft, the area to target is manufacturing.

Manufacturing costs can be reduced dramatically through careful planning of material acquisition and tooling time. Figurel0.2.2 shows a breakdown of the factors affecting the manufacturing costs. Personnel costs dominate 82% of the manufacturing expenses, thus comprising 64% of the total aircraft cost. Waste disposal comprises 11% of the manufacturing cost or almost 9% of the total cost of the aircraft. Tooling comprises 7% of the manufacturing costs and only 5% of the total aircraft cost. Thus the top three dominating factors affecting the total cost of the aircraft were the personnel (64%), the subsystems (16%), and the waste disposal (9%). The costs of the subsystems were fixed by the design. The personnel and waste disposal costs depend on the efficiency of the manufacturing process.

Table 10.2.1: Cost Breakdown of The Elite $ 10.00 Avionics Battery Pack Switch Harness $ 5.00 Miniature Servos (3 @ $35) $105.00 $ 5O.00 Electric Motor Speed Controller Astro- 15 Motor $107.00 Batteries (13 @ $4) $ 52.00 $ 4.00 Motor Power Wiring (2 feet) $ 10.00 Landing Gear Zin ali 10-8 $ 9.58 $ 462.58 Raw Materials: $160.00 Manufacturinl[: $1800.00 Personnel Costs (180 hrs @ $10/hr) Tooling $150.00 Waste (1.5 lbs @ $10/oz) $ 240.00 $2190.00 s+Manufacturin $2812.58 Figure 10.2.1: Breakdown of Total Aircraft Cost (16%) Subsystems (6%) Raw Materials Manufacturin(g 78%) Figure 10.2.2: Breakdown of Manufacturing Costs Waste Disposal (11%) (7%) Tooling Personnel I82%) The effects of a slight decrease in either the estimated number of man-hours or the amount of waste disposal on the total cost of the aircraft is illustrated in Figure 10.2.3. A

decrease of 20 man-hours decreases the total aircraft costby approximately $314. Thus

for everyman-hour,thetotal aircraftcostincreases by $15.68. Notethatthis numberwas

greaterthanthe $10/hrratepaybecause of the overhead factorandthe profit allowance.

The effectsof decreasing theamountof wastedisposalwasnot aspronounced.A

1 Ib decrease in wastedisposaldecreases thetotal costof the aircraftby $250. A

decrease any lessthan 1 lb is improbablebecause theinitial wastedisposalestimatewas

only 1.5lbs.

Figure 10.2.3: Influence of Waste Disposal and Production Hours on Cost

5500 "_ .dkJL

ooo

Waste Disposal 4500 _____m < 2.5 lbs 1.5 lbs 4000 _¢_j_l_ Q_ 4...* o) C3 -II-- C3 1.0 lbs o F-- r" 1 O0 120 140 160 180 200 220 240 260 Number of Man-Hours 10.3 Economic Performance Some cost measures of merit were also calculated for The Elite. They can be found in Table 10.3.1. These merit factors were based on depreciation costs, maintenance-insurance costs, and fuel costs. Depreciation costs recognize that the flight hardware has a limited life. Current technology in Aeroworld permits a total lifetime of 100 flight hours. Also, depreciation costs increase with range and decrease with cruise speed. The depreciation costs for The Elite were $5.10/flight. The Maintenance- Insurance costs increase with the design speed (60 ft/sec) and the maximum takeoff weight (4.88 lbs) and for The Elite were $0.20/flight. The fuel costs per flight depend on the current draw, the flight time (0.1157 hour), and the fuel cost ($3.08/amphour) and for The Elite were equal to $4.65/flight.

IO0 Table 10.3.1: Economic Figures of Merit for The Elite $5.10 Depreciation Cost/flight $ 0.20 Maintenance-Insurance Cost/flight $ 4.65 Cost per Flight (CPF) $ 9.95 $1.43 Cost per Flight Minute (CPFM) $ 0.335 Cost per 1000 feet (CP1000) The cost per flight (CPF) is the sum of the depreciation costs, the maintenance- insurance costs, and the fuel costs. The depreciation costs are high because they are directly related to the total cost per aircraft, which for The Elite is fairly high. However, if the total cost per aircraft proves to be high even after manufacturing is completed, the depreciation costs could still be decreased by decreasing the flight time. The other major contributor to the CPF is the fuel costs which also could be decreased by decreasing the flight time. For instance, keeping the cruise speed at 60 ft/s but flying only 15000 ft instead of the entire range of 25000 ft would decrease the CPF by 62% to only $6.12/flight.

The cost per flight minute (CPFM) is directly proportional to the CPF and inversely proportional to the flight time. Moreover, since the CPF is directly proportional to the fright time, the CPFM is actually independent of the flight time. Similarly, the cost per 1000 feet (CP1000) is directly proportional to the CPF and inversely proportional to the range. These dependencies correspond to the CP1000 being inversely proportional to the cruise speed. Thus, a decrease in flight time by decreasing the cruise speed would decrease the cost per flight, increase the cost per 1000 feet, and keep the cost per flight minute the same. Thus, if the pilot desires to minimize the cost per flight, he could do so by decreasing the distance of the flight, keeping the cruise speed and thus the CPFM and CP1000 the same. To minimize the cost per 1000 feet (thereby increasing the cruise speed), a trade off must be made between the cost per flight and the cost per 1000 feet.

APPENDIX A

Critical Design Summary: 2-17-94 2-24-94 3-3-94 Final Design

Parameter ] Initialsol j 2-10-94 i

Rh !

*(all distances relative I to aircraft nose and in common units]* I DESIGN GOALS: V cruise Tuan 60 ft/sec 65ft/sec 65ftJsec 65 Wsec 160Wsec 6 6 6 6 No. of passengers/crew 1 Doug 6 Max Range at Wmax Tuan 30,500 ft 30,500 ft 30,500 ft 30,500 ff 32336 11 Max RIC at SL Steve 16.1 rids Altitude cruise Doug 50 (11) 50 ft Minimum turn radius Doug 20 ft Max Range at Wmin I Tuan 33000 ft 4.25 Maximum TO Welght-WMTO (Ib) 4.25 4.25 Doug/Dan 4.1 4.1 4.1 Wing Area 6.5 if^2 6.5 ft^2 6.5 6.5 ft^2 Maximum TO Weight - WMTO (Ib) DouoJDan 4.22 Ib 4.22 Ib 4.88 q4.88 Ib Empty Flight Weight Doug/Dan 4.17 Ib 4.17 Ib 4.83 !4.83 Ib Wing Ioading(WMTO) Ib/ft^2 Jonathan .65 Ib/scl.ft. i10.3 oz/ft^2 0.75 0.75 Ib/ft^2 max length (in) Jonathan 38 inches 38 inches :.._ _ :_ ......... 38 inches 'nax span (ft) Amy 7.21 ft 8 ft 7.66 ft "nax height Jonathan 1.5 ft Fetal Wetted Area Jonathan 17.0 sq ft 17.0 SOl ft WING _,spect Ratio Amy 8 9 9 ._pan (incuding fuselage) Amy 7.2111 8 11 7.66 ft J, rea (ft^2) Amy 6.5 6.5 6.5 6.5 ft^2 =loot Chord Amy !0.9 ft 10 3/16 inches 10 3/16 inches rip Chord Amy 0.9ft 10 3/16 inches 10 3/16 inches :aper Ratio Amy _1

] 1

3 mac - MAC Jonathan -0,0582 -.0.0582 eading edge Sweep Jonathan 0 Jonathan 0 1/4 chord Sweep *

o

)ihedral Jonathan 8 degrees 5 degrees rwist (washout) Amy 0 DF101 _.irfoil section Jonathan ]DF101 )esign Reynolds number Jonathan _350,000 350.000 Jc Jonathan 0.11 0.11 ncidence angle (root) Jonathan 0.72 degrees 1.2 degrees Hor. poa of 114 MAC Jonathan 13.03inches 13.03 inches Ver. poa of 114 MAC Jonathan e- Oswald efficiency 0.949 Amy 0.949 J Amy 0.056 0.008 CDo -wing CLo - wing 10.0995 0.0095 CLalpha -wing 0.083/degree .083/degree FUSELAGE Length [ Amy 38 inches 38 inches 38 inches 38 inches Cross section shape Amy circular 5 in dia r__;_ _:_: ¢: circular- 5 in diam.

NominalCrossSectionArea (in^2) Amy (average diameter = 4 [ 12.57 0.0873 11^2

,n) [

Finesa ratio Amy i7.6 7.6 Payload volume (cubic in.)

Amy 25 in_3 25 in'_3 25 ir_3 25 in_3 Planform area Amy 1.04 1.06 ft^2 Frontal area (sq. ft.) Steve 0.155 0.155 0.155 0.3595 0.155 sq. ft.

CDo - fuselage Steve 10.0194 0.0194 0.00608385 0.00608385 CLalpha - fuselage _ __ EMPENNAGE I ! :, Horizontal tail [ [ i :

_rea (ft^2) Dan

;1.2 il .6 0.9 09 ft^2 span Dan il ,8 2.4 1.8 1,8 ft aspect ratio Dan 9 3.6 3.6 3.6 rOOt chord Dan I 18 inches 0.5 ft 0.5 ft lip chord Dan !8 inches 0.5 ft 10.5 ft average chord Dan 18 inches 0.5 ft 0.5 ft taper ratio Dan 1

i

Le. sweep Dan 0 15 degrees 15 degrees 15 degrees 1/4 chord sweep Dan 0 15 degrees 15 degrees 15 degrees incidence angle Dan 0 -1.32 -2 degrees deg. rees hor. pos. of 1/4 MAC Ban 34.0 inches 34.5 inches vet. pos. of 1/4 MAC Ban 10.4 in. above ground A.irfoil section Jonathan SD8020 =,- Oswald efficiency Jonathan 0.956 0.956 _Do -horizontal Steve 0.0014 0.00110769 !0.00110769 CLo-horizontal - Jonathan 0 0 0 3Lalpha - horizontal Dan 3.803

3.803

3M mac - horizontal Jonathan

0 i0

Vertical Tail ;rea (ft^2) Amy 0.55 0.3 if^2 2.2 11.2 _.spect Ratio Amy 'dOt chord

o.5_ )o5_ Dan

:ip chord Dan 0.5 ft 0.5 ft _verage chord Dan 0.5 ft _0.5 ft :aper ratio Dan 1 .e. sweep Dan 45 degrees 1/4 chord sweep Dan ;49 degrees hor. poa. of 1/4 MAC Dan 34 34.5 in vert. poa. of 1/4 MAC Dan 14.875 degrees Airfoil section Jonathan flat plate flat plate flat plate flat plate SUMMARY AERODYNAMICS Jonathan

CI max (airfoil) 1.14 I¸_ ii_i_iiii!_iiii_iiiiilzi!iiiii_i_ii 1.14

Cmo (airfoil) Jonathan -0.0582 -0.0582 CL max (aircraft) Jonathan 1.03 1.03 lift curve slope (aircraft) Jonathan 0.083 deg-1 0.083/degree Steve CDo (aircraft) 0.1282 0.0289 0.035

_ ili!,ii_iiiiiiiii!!!i!iii_iiiiii,iii!i!iiii_ii 0.0325

Steve efficiency - e (aircraft) 0.733 0.89 0.89 Jonathan 12.333 ° Alpha stall (aircraft) 12.333 degrees Alpha zero lift (aircraft) ° Jonathan -0.23 -0.23 L/D max (aircraft) Jonathan 14.8 14.8 Alpha LK) max (aircraft) Jonalhan 9.0 ° 9.0 degrees WEIGHTS (pounds) 4.17 Ib 4.173 Weight total (empty) Doug/Oan 4.834 4.834 C.G. most forward-x&y Doug 13.36 13.36 C.G. most aft- x&y Doug 13.471 i13.471 Avionic= (Ib)

bow 0.31 10.31

!0.14

Ooug Control Linkages (Ib) 0.14

Payload-Pasa.&lugg.-max (Ib) Douoj'Dan 0.74 oz !0.74 oz 0.0463 !0.0463 i 12 oz 12 oz 0.657 0.657 Encjine & Engine Controls (Ib)

Exw

Propeller (Ib) Tuan O.0741bs 0.097 0.097 Tuan Fuel (battery) (Ib) 1.38 Ibs 1.38 tl.38 Structure Jonathan Wing (Ib) 1.05 1.05 Jonathan Fuselage (Ib) 0.5 10.5 Dan Main Landing gear (Ib) 0.256 10.256 Nose gear (Ib) Dan 0.173 10.173

empenaga weight (Ib) Dan O. 15625 to. 15625

PROPULSION i Type of engines Steve Astro 15 Astro15 number Tuan 1 ;;:_; _ ....

)lacement Steve nose (tractor) nose (tractor) Pavail max at cruise (60 ft/s) Steve/Tuan 1170 watts r 170 watts 132 watts Preq cruise Steve/l'uan 65 watts I 65 watts 77.6 watts max. current draw at TO Tuan _ ,!16 amps 14.292 A 4.2 amps

:ruise current draw Steve

17.78 A 110 amps 110,357 amps 13,5A _ropeller type Tuan Zinger 11- iZinger 10-7 !John Brothers11- _Zingati 10-8 Z ngali 10-8

7 16

"ropeller diameter Tuan 110" 10" 10in _ropeller pitch Tuan 7 8 8 j8 in _lumber of blades Tuan 3 3 !3 • nax. prop. rpm Steve/Tuan ; 16310 rpm :ruise prop. rpm Steve/'l'uan i i ,13687 rpm nax. thrust Steve/'r'uan i 4,653 Ib 14.3 Ibs :ruise thrust Stevefl'uan !1.05 Ib Steve/'l'uan )attery type P-130SCR P-130SCR !P-130SCR number Steve/Tuan 13 13 113 individual capacity Steve/Tuan 1300 mah 1300 mah "1300 mah 1.2 individual voltage Steve/Tuan 1.2 il.2 Steve/l"uan !1300 mah pack capacity 1300mah 1300mah pack voltage Steve/Tuan 15.6 15,6 STAB AND CONTROL : i i i Neutral point Dan i41% MAC Static margin %MAC Dan ; i 111% MAC O_u'l Hor. tail volume ratio !0.359 1 0.359 r03_ l i Vert. tail volume ratio !

Amy 0.0105 Dan Stabilator area (ft^2) 0.9 0.9 ft^2 I i Stabilator max deflection Dan i 120 deg (tail) 12 ° 112 degrees Rudder Area Dan 0.9 i0.3 if^2 Rudder max deflection Dan 15 ° [15 degrees Oal3 Aileron Area (ft^2) 0.75 !0.75 ft^2 Dan Aileron max deflection i i15° 15 ° 115 degrees ! 0.655 Cm alpha 0.655 i-0.45 Cn beta i -1.1 -1.1 i0.025 CI alpha tail 0.0663 0,0663 _1 delta • tail i 0.0663 0.0663 PERFORMANCE

f

r

Vmin at WMTO Tuan i24,8 ft/s

Vmax at WMTO Steve r775 77.5 !71.71 Ws

_/stall at WMTO Jonathan I 23,44 23.44 24.8 ft/s Range max at WMTO Tuan !32919.3 feet Endurance @ Rmax Tuan I 14,86 min Endurance Max at WMTO Tuan i ! 118.62 min Range at @Emax Tuan ] ' !27933ft [ Range max at Wmin Tuan i !33790 ft @ V = 35 ROC max at WMTO Steve I i ;16 fl/s @ V = 36.5 Ift/s _.bs. Ceiling Steve ,! il 94306 ft _in Glide angle Steve ; 3.87 = rio distance at WMTO Steve 1 126.0 ft Steve Route Percent Servicable at Cruise SYSTEMS

! I !

Landing gear type

Do_ tricycle

tricycle

_lain gear position Doug

15.5 in

Vlain gear length Doug 5 in

2" diameter _lain gear tire size

C_g

2 in diameter

_ose/tail gear position Ooug

4.375" beh prop 14.375 in behind prop 1/t gear length Doug i 15 in 1/t gear tire size Doug I 2 in diameter _ngine speed control Doug TEKIN ! TEK N speed control Control surfaces Doug stabilator, rudder i TECH DEMO - Final I Max Take-Off Weight

oow

Empty Operating Weight

Doug

Jonathan Ning Area 6.5 _2 Hor. Tail Area Dan Vert Tail Area Dan _,.G. position at WMTO Doug i I !

I/4 MAC position Jonathan 13.03 in 3tatic margin %MAC Dan i r / takeoff J Steve 24.8 fl/s i [24.8 ft/s _ange max I Tuan

_.irframe struct, weight : Doug

Tuan )ropulsion sya. weight (Ib) 2.134 2.134 Ib _,vionics weight

Ooug

.anding gear weight

: Doug

ECONOMICS: • aw materials Cost Amy ;150 $160.00 !$150 )ropulsion system cost Amy 150 _$172.58 j150 zvionica system cost i Amy i150 150 !$280.00 )roduction manhours 180 PlSOhours i Amy j180 )ersonnel costs $1,800.00

i1800

Amy :ooling costs 500 $150.00 i Amy 1500 :otal cost per aircraft 14312 $4,394.45 Amy i4312 current draw at cruise WMTO Amy r17.78 A 13.5 A 10 amps 10.233 amps I

r

CPF at Vcruise and Rmax ! Amy $8.37 ! i : $1.21 CPFM i Amy CP1000 i Amy

$0.34

APPENDIX C:

Computer Codes

C C C Computer code to determine the horizontal tail areas C and locations needed to rotate at take-off at certain C tail deflection angles.

C REAL dLqft,dtail,Lifl,W,ltt,lt FORMAT(8x,f8.4,8x,f5.3,3x,f5.3,5x,f6.2,8x,f6.3,4x,f8.4) C OPEN (UNIT=2, FILE='fl ') C OPEN (UNIT=3, FILE='f2') OPEN (UNIT=4, FILE='f3') C OPEN (UNIT=5, FILE='f4') Pi = 4.0*ATAN( 1.0) W=4.8 rho = 0.002378 Vto = 25.0 cw = .84896 bw = 7.646 Sw = 6.49 ARw = bw**2/Sw alplo = 2.0"Pi/180.

Clalpw = 5.73 CbLalpw = Clalpw/(1.+Clalpw/(Pi*Arw)) CLo = .0995 WRITE(6,*) Clalpw,CbLalpw, CLo Lift = 0.5*rho*CLo*Sw*Vto**2 dWeight = 2.25/12.0 dLift = 4.0228/12.0 ct = 6.0/12.

C St = 2.0 C alpt = - 15.*Pi/180.

It = 21./12.

tLifl morn' WRITE(6,*) ' alpha tail (deg) ct St It DO 20 alptt = 8.,20.,2 WRITE(4,*) ' ' WRITE(6,*) ' ' alpt = -alptt * Pi/180.

C DO 5 ctt = 5.0,8.5,.2 C ct = ctt/12.

DO 10 St - .5,2.5,.01 DO 30 ltt - 16.,24.,.01 It =ltt/ 12.

bt = St/ct ARt = bt**2/St C ARt = 3.6 dtail = It - dWeight Clalpt = 5.73 CbLalpt = Clalpt/(1.+Clalpt/(Pi*ARt)) tLift = 0.5*rho*CbLalpt*alpt*St*Vto**2 rmom = Lifl*dLifl - W*dWeight + ABS(tLifl)*dtail C WRITE(6,33) alpt* 180/Pi,ct* 12.,St,It* 12.,tLift,rmom IF (ABS(rmom) .LT. 6.0E-5) THEN WRITE(4,*) ltt, St C WRITE(6,*) '*******', alptt,Ltt, St ENDIF C WRITE(2,*) alptt,rmom C WRITE(3,*) ctt, rmom C WRITE(4,*) St, rmom C WRITE(5,*) ltt, rmom 30 CONTINUE CONTINUE C WRITE(4,*) ' ' C WRITE(5,*) ' ' CONTINUE C5 20 CONTINUE C CLOSE (UNIT=2) C CLOSE (UNIT=3) CLOSE (UNIT=4) C CLOSE (UNIT=5) STOP END PROGRAM AE350 C PROGRAM TO AUTOMATE LIFTING-LINE THEORY C C Declare variables and types C REAL lambda, AR, twist, AROOT, pi REAL cam(50), THETAP, mo , Y(50) REAL THETAN(50), YDISTN(50), CRDN(50), MON(50) REAL DI, D2, M(50,51), A(50) REAL CL(50), L(50), COSTH(50) REAL SIGMA, CLW, CDI INTEGER SPAN, I, J, K, astate, sstate *****************WW*W****W*WWWW*****WWWWWW**WWW**W*WWW*WWWWWWWWWWW*********** C C Prompt user for necessary input and program parameters/open datafiles C ***************************************************************************** SPAN = 10 pi = 4.0*atan(l.0) lambda = 1.0 write(0,*) 'Enter the Aspect Ratio' Read(5,*) AR twist = 0.0 twist = twist*pi/180.0 Do 4 I -- i, SPAN THETAN(I) = pi*I/(2.*SPAN) YDISTN(I) -- COS(THETAN(I)) write(6,*) 'Enter the max camber at the root.'

read(5,*) cam(l) cam(l) = cam(l) "3.1415926/180.0 Do 6 I= i, SPAN cam(I) = cam(l) AROOT = 0"3.14159/180.0 mo = 5.73 do 43 I = I, SPAN MON(I) = mo 43 continue rho = 0.00238 visnu = 0.00015723 Write(6,*) 'Enter the flight velocity' Read(5,*) vinf Write(6,*) 'Enter the planform area' Read(5,*) Sarea Write(6,*) 'Enter the CLmax of the airfoil' Read(5,*) CLmax Sstep = 0.01 Astep = 1.0"3.1415926/180.0 sstate = 0 astate = 0 CLsuml = 0.0

C

C Begin program routines for calculating coefficient matrices

C

C

109 Continue

do 7 I = i, SPAN

1.O --

CRDN (I) =

(I. 0-1ambda) *COS (THETAN (I))

continue

do 30 I = i, SPAN D1 s MON(SPAN)/ (CRDN(I) *MON(I) ) D2 = MON(SPAN)/(2.0*AR*(I.+ Iambda)*SIN(THETAN(I))) do I0 J-- 1,SPAN K-- 2*J - 1 M(I,J) = (DI + D2*K)*SIN(K*THETAN(I)) continue 3O continue ******************************************************************** C C Find absolute angles of attack (left side of the equations) C do Ii0 I = A(I ii0 CL( Tlift = 0.0 SIGMA = 0.0 CLsum = 0.0 do 69 I = 1,SPAN M(I,SPAN+I) = AROOT - twist*COS(THETAN(I)) + cam(I)*(I.0-COS(THETAN(I))) C WRITE(6,*) M(I,N+I) 69 continue C C ******************************************************************** C C Send augmented matrix M(SPAN, SPAN+I) to subroutine to solve C for the coefficients Ai through Aspan by gaussian elimination C ********************************************************************* CALL SIMEQN (M, A, SPAN) C ********************************************************************** C C Open data files to prepare for output and plotting of results C ********************************************************************** C OPEN(UNIT=69, FILE = 'CLout') OPEN(UNIT=79, FILE = 'Lout') dimen = 0.5*rho*vinf*vinf*Sqrt(Sarea/AR) b = Sqrt(AR*Sarea) c = b/AR do 16 I = i, 50 MON(I) = mo THETAP = 3.141592"I/(2.0"50.0)

do 12 J _ I, SPAN

K = 2*J - 1

CL(I) = CL(I) + A(J)*SIN(K*THETAP)

12 continue

C

C

C Calculate sectional lift coefficient (CL) and Lift per Span

C

CL(I) = CL(I)*MON(SPAN)/(I.0 - (I.0-1ambda)*COS(THETAP))

If (CL(I) .GT. CLmax) CL(I) -- 2.0*CLmax - CL(I)

L(I) = CL(I)*(1.0 - (l.0-1ambda)*COS(THETAP))*dimen

Y(I) -- COS(THETAP)*b/2.0

write(69,*) Y(I), CL(I)

COSTH(I) = COS(THETAP)

continue

C

C Lift at center line is equal to Cls*SIN(pi/2) -> Cls

C do 3 I = I, 49 CLsum = CLsum + (CL (I) +CL (I+l))/2.0* (Y (I) -Y (I+l)) Tlift = Tlift + (L(I) + L(I+I))*(Y(I)-Y(I+I)) do 57 J = 1,50 write (79,*) Y(J), L(J) 57 continue C C C Determination of Oswald efficiency factor (SIGMA) C do 85 I = 2,SPAN K = 2"I - 1 SIGMA = SIGMA + K*A(I)**2/(A(1)**2) 85 continue write(6,*) 'Correction factor (sigma) =' , SIGMA write(6,*) 'Oswald efficiency factor for the wing:', 1.0/(I.0 + SIGMA) C C C Calculation of wing lift coefficient (CLW) and induced drag coeff. (CDI) C C CLW = MON(SPAN)*pi*A(1)/(2.0 + 2.0*la_da) CDI = (CLW**2/(pi*AR)) * (I. + SIGMA) Write (6, * ) 'Wing Lift Coefficient at AOA:',_OOT*I80/pi,' (deg)',CLW Write (6, *) 'Necessary Wing area', Sarea write (6, * ) write (6, *) 'Induced drag coeff for the wing:', CDI write (6, *) write (6,*) 'Total Integrated Wing lift', Tlift write (6, *) write (6, *) 'Center Line Section CL', CL(50) write (6, *) write (6, *) 'Reynolds nu_er =', c*vinf/visnu weight = 3*4.76 write(6,*) ' ' write(6,*) 'weight', weight

write(6,*) 'span', b

write(6,*) 'cord', c

write(6,*) 'vinf', vinf

write(6,*) 'CLsum', CLsum

CLOSE (69)

CLOSE (79)

IF ((ABS(weight-Tlift) .GT. 0.0001) .OR.

(ABS(CLsum-CLsuml) .GT. 0.0001)) Then

IF (weight .GT. Tlift) Then

IF (sstate .EQ. i) Sstep =-Sstep/2.0

sstate = 0 Else IF (sstate .EQ. 0) Sstep = -Sstep/2.0 sstate = 1 Endif vinf _ vinf + Sstep IF (CLsum .GT. CLsuml) Then IF (astate .EQ. I) Astep = -Astep/2.0 astate = 0 Else IF (astate .EQ. 0) Astep = -Astep/2.0 astate = 1 Endif AROOT = AROOT + Astep CLsuml = CLsum GOTO 109 Endif stop end SUBROUTINE SIMEQN(A, X, N) C This program solves a set of simultaneous equations by Gaussian C elmination to create an upper triangular system, followed by C back-substitution to abtain the solution.

C C Maximum number of equations in set.

C INTEGER MAXEQN PARAMETER (MAXEQN=50, MAX2=MAXEQN* (MAXEQN+I)) C C Local variabiiles C REAL A(MAXEQN, MAXEQN+I), X(MAXEQN), DET, LARGE INTEGER N, I, J LARGE = 0.0 DET = 1.0 C C C Read in number of equations C C Check to see if maximum equations is exceeded.

C DO I0 I = I, N DO 13 J = I,N+I CONTINUE i0 CONTINUE IF (N .GT. MAXEQN) THEN WRITE (0,*) 'Maximum number of equations is exceeded.'

WRITE (0,*) '(Maximum number=',MAXEQN,')' STOP END IF Read in coefficients and constant into the arrays and find

C

largest value

C

C

DO 99 I -- I,N

DO 88 J-- I,N+I

IF (ABS(A (I, J) ) .GT. LARGE) LARGE=ABS(A(I,J))

88 CONTINUE

99 CONTINUE

C

C Use Gaussian elimination to triangulate the equations

C

CALL GAUSS(A, MAXEQN,N, LARGE)

WRITE(6,*) 'LARGE:', LARGE

Solve new set by back-substitution

CALL BACK(A,MAXEQN, N,X, DET)

C WRITE(6,*) 'DET:', DET

C

C Calculate and print answer confidence ratio

C

IF (ABS(DET/LARGE).LT. 2E-7) THEN

WRITE(6,*) 'WARNINGDETERMINANT RATIO IS SMALL!'

END IF

C

Print results

C

C

C WRITE (6,101) ABS(DET/LARGE),DATSET, (I, X(I), I-I,N)

C I01 FORMAT(/,'Determinant Ratio=',Fl0.5,//,'The solutions to the data

C + set,',X,Al2, 'is:',//,3('A(',I2,') = ',EI3.5,TR5),/,)

C

RETURN END

C

Subroutines Gauss and Back follow below

C

C

C

SUBROUTINE GAUSS(A, MAX, N, LARGE)

C

C This subroutine triangulates a set of N equations Ax=B

C

using Gaussian elimination.

C

C Dummy arguments (temporary since actual arguments are in MAIN

C

INTEGER MAX,N REAL A(MAX,MAX+I),LARGE

C

Local variables

C

C

REAL PIVOT,ABSP,MULT,TEMP INTEGER PROW, I,J,K Eliminate one element from each row in turn DO 40 I=I,N

C

Set initial values for PIVOT and PROW

C

C

PIVOT=A (I, I) ABSP=ABS (PIVOT) P ROW= I

C

C Look for a larger potential pivot

C

DO 10, J=I+I,N IF (ABS(A(J,I)).GT.ABSP) THEN P IVOT=A (J, I ) ABSP=ABS (P IVOT ) PROW-J END IF

i0

CONTINUE Check to see if pivotal element is "zero" IF (ABS(PIVOT/LARGE) .LT. 2E-7) THEN WRITE(0,*) 'Ill-conditioned system! Process aborted!'

STOP END IF Was a larger pivot row found?

IF (PROW.GT.I) THEN DO 20, J--I,N+I TEMP=A (I, J) A (I, J) =A (PROW, J) A (PROW, J) =TEMP

20 CONTINUE

END IF Eliminate coefficients of X(I) from rows I+l to N DO 30 J=I+I,N MULT=A(J,I)/PIVOT A(J,I)=0.0 DO 25 K=I+I,N+I A (J, K) =A (J, K) -MULT*A (I, K) 25 CONTINUE 30 CONTINUE 40 CONTINUE C C Gaussian elimination is completed C RETURN END SUBROUTINE BACK(A,MAX,N,X,DET) C C This subroutine solves an upper triangular system of simultaneous C equations through back-substitution.

C C Temporary arguments C INTEGER MAX,N REAL A(MAX,MAX+I),X(N),DET C C Local variables C REAL QUOT INTEGER I,J DO 77 I= I,N DO 66 J= I,N+I C WRITE (6, *) A(I, J) 66 CONTINUE CONTINUE C C Calculate X(N) C X (N) =A (N, N+I)/A (N,N) C C Now calculate remaining values in reverse order C

DO 20 I--N-l,l,-I

QUOT=A (I, N+I)

DO i0 J=I+I,N

QUOT=QUOT-A (I, J) *X (J)

I0 CONTINUE

X (I)=QUOT/A (I, I)

2O CONTINUE

Evaluate the determinant

DO 50 I=I,N

DET=DET*A (I, I )

A(I,I)

C WRITE(6,*) 'A(',I,I,') :',

C WRITE (6, *) ' DET*: ',DET

5O CONTINUE

C

complete

C

Solution by back-substitution method

C

RETURN

END

C RPV Propulsion Program C bv Tuan A. Le C March 24, 1994 C Design Program to calculate the power available, power required, C current draw, motor rpm, range, and endurance for a propulsion system.

REAL n,Ia,Kt,Kv,Nm,Np,Ng,J,Mah,Wgt(3),Volt(5) OPEN(UNIT= 10,FILE='POWER',STATUS='UNKNOWN') OPEN(UNIT= 11 ,FILE= 'CURRENT' ,STATUS= 'UNKNOWN') open(12,file='PAV') write(11,*) 'Velocity Range Endurance Ia Vact J Preq Pavail Nm etap' Mah = 1400.

batcap = Mah/1000.

PI = 4.*atan(1.)

rho = 0.002378 CDo = 0.0325 Clmax = 1.03 AR = 9.0 S = 6.5 b = sqrt(AR*S) c=b/AR e = 0.89 n = 1.0 Wgt(1) = 4.83-.55 Wgt(2) = 4.83+.05 Wgt(3) = 4.83+.55 Volt(l) = 1.4 Volt(2) = 8.0 Volt(3) = 10.5 Volt(4) = 12.8 Volt(5) = 15.6 Kv = 7.8568E-4 Kt = 1.097846 Tloss= 1.372935 Tloss = 0.

Ra = 0.120 Rbat = 0.08 Ng = 0.95 Ng= 1.

Dprop = 10./12.

GR =31./14.

C Code to calculate Pavail and Preq

do 2000ii = 2,2

W = Wgt(ii)

VINIT=sqrt(2.*W / (RHO*S*CLMAX))

do 1000 Vact = 6.0,15.7,.1 c do 1000 iii = 1,5 idone = 0 c Vact = Volt(iii) do 30 V=Vinit,75.,0.01 delta = 500.

Nm = 6000.

Q = .5*rho*V**2 CL = n*W/(Q*S) CD = CDo+CL**2 / (PI*e*AR) Treq = 1.356*CD*Q*S Preq = 1.356*0.5*rho*V**3*S*CD 2O CONTINUE Ia=(Vact-Kv*Nm)/(Ra+Rbat) PmoA=7.397E-4*(Kt*Ia-Tloss)*Nm Np=Nm/GR J=V*60./(Np*Dprop) C Zingali Cp and eta curve fits Cp=-3.460082E-l*J**3+4.064685E-l"J'*2-2.062646E-1*J + 1.305385E-1 etap= -6.734308E+0*J**3+1.143146E+ 1*J**2-5.931978*J+ 1.662525 C Graupner Cp and eta curve fits c Cp=-2.253302E-1*J**2+1.181663E-1*J + 8.265346E-2 c etap=-5.002783*J**3+5.884037*J**2-1.799132*J+Z734548E-1 PmoB=1.356*Cp*rho*(Np / 60.)**3*Dprop**5 / Ng if(PmoA .LT. 0.) then print *, 'PmoA negative' goto 999 endif if(PmoB .GT. PmoA) then delta = delta/2.

Nm = Nm-delta goto 20 endif if(abs(PmoA-PmoB).GT.1.E-3) THEN Nm=Nm+delta goto 20 endif if(J .GT..75) then print *, 'Advanvce Ratio too high' goto 999 endif

Pavail=7.397E-4*(Kt*Ia-Tloss)*Nm*Ng*ETAp

if(Pavail .LT. 0.) then

print *, 'Pavail Negative'

goto 999

endif

c write(10,*) V,Preq,Pavail,Vact,J,Nm,Np

C Calculate Rangeand Endurance

ENDURANCE = (batcap) / Ia*60.

RANGE = ENDURANCE*V*60.

if((idone .EQ. 0) .AND. Pavail .LT. Preq) then print *, 'file: V,RANGE,ENDURANCE,Ia,Vact,W',V,RANGE,ENDURANCE,Ia,Vact,W write(11,4444) V,RANGE,ENDURANCE,Ia,Vact,J,Preq,Pavail,Nm,W print *,'V,J,Preq,Pavail',V,J,Preq,Pavad idone = 1 goto 30 endif 30 continue 999 continue idone = 0 print * write(10,*) 1000 continue print * write(11,*) 2000 continue 4444 FOICMAT(F6.2,F10.2,F8.3,F8.3,F7.3,F9.5,F8.3,F7.2,F10.2,F6.3) STOP END

APPENDIX D

Zingali 10-8 Propeller Data INPUT TO PROP123 A.) Propeller Designation: Zingalil0-8 B.) Number of Blades: 3 Diameter: 10.00000 C.) Airfoil section selected: NACA44XX Low RE D.) Blade thickness entered as: INCHES E.) Blade data entered at radial locations specified as: INCHES F.) Radius at which blade setting is measured: 4.00000 G.) Blade setting (i.e. ref angle for whole blade): 17.6570 H.) Number of radial data positions (3-9): 9 I.) Data Point Radius* Chord Thickness* Angle 1 1.000 0.723 0.337 51.854 2 1.500 0.802 0.271 40.325 3 2.000 0.861 0.230 32.482 4 2.500 0.889 0.206 26.990 5 3.000 0.862 0.171 22.997 6 3.500 0.785 0.147 19.991 7 4.000 0.686 0.113 17.657 8 4.500 0.577 0.096 15.798 9 5.000 0.429 0.081 14.287 J.) Refinement Analysis: ANALYSIS INCLUDING INDUCED VELOCITY AND TIP LOSSES K.) C1/Cd coefficient adjustments: MACH AND REYNOLDS NUMBER ADJUSTMENTS L.) Altitude: 0. feet Rho: 2.37690E-03 M.) Airspeed FIXED at: 40.9000 N.) Range of Advance Ratio to be used in calculations: Jmin: 0.340000 Jmax: 0.800000 OUTPUT FROM PROP123 FOR ZINGALI 10-8 PROPELLER PERFORMANCE ESTIMATE for zingali 10-8 Analysis options: RA= 3 and LDA = 4 Fractional rad, X: 0.30 0.45 0.60 0.70 0.75 0.80 0.85 0.90 0.95 Radial position, r: 1.50 2.25 3.00 3.50 3.75 4.00 4.25 4.50 4.75 Blade chord, C: 0.80 0.88 0.86 0.79 0.74 0.69 0.63 0.58 0.51 Thickness, In: 0.12 0.13 0.13 0.12 0.11 0.10 0.10 0.09 0.08 Thickness ratio, T: 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 0.15 Blade Angle, Beta: 40.33 29.52 23.00 19.99 18.75 17.66 16.67 15.80 15.02 GeomewicPitch,GP: 8.00 8.01 8.00 8.00 8.00 8.00 8.00 8.00 8.01 Solidity, S: 0.153 0.168 0.165 0.150 0.141 0.131 0.121 0.110 0.097 THRUST, POWER, EFFICIENCY, AND VELOCITIES J: 0.400 0.430 0.460 0.490 0.520 0.550 0.580 0.610 0.640 0.670 0.700 Ct: 0.149 0.139 0.132 0.127 0.120 0.114 0.107 0.101 0.094 0.087 0.078 Cp: 0.089 0.088 0.087 0.086 0.084 0.083 0.080 0.078 0.075 0.072 0.067 eta: 0.667 0.676 0.697 0.720 0.738 0.756 0.773 0.788 0.800 0.809 0.813 Mt: 0.462 0.430 0.403 0.378 0.357 0.338 0.321 0.306 0.292 0.280 0.269 RPM: 11700 10884 10174 9551 9000 8509 8069 7672 7312 6985 6685 Thrust Distribution: (dCt/dX vs. X and J) X J:0.40 0.43 0.46 0.49 0.52 0.55 0.58 0.61 0.64 0.67 0.70 0.30 0.07 0.07 0.06 0.06 0.05 0.05 0.05 0.05 0.04 0.04 0.04 0.45 0.17 0.13 0.12 0.12 0.11 0.11 0.10 0.10 0.09 0.08 0.07 0.60 0.21 0.21 0.19 0.19 0.18 0.17 0.16 0.15 0.14 0.13 0.11 0.70 0.25 0.24 0.23 0.22 0.21 0.20 0.19 0.17 0.17 0.15 0.14 0.75 0.26 0.25 0.24 0.24 0.22 0.21 0.20 0.18 0.18 0.16 0.14 0.80 0.27 0.26 0.26 0.24 0.23 0.22 0.20 0.20 0.18 0.17 0.15 0.85 0.27 0.26 0.26 0.25 0.23 0.22 0.21 0.20 0.18 0.17 0.15 0.90 0.27 0.27 0.26 0.24 0.23 0.22 0.21 0.20 0.18 0.17 0.15 0.95 0.37 0.35 0.33 0.31 0.30 0.28 0.27 0.25 0.23 0.22 0.20 Torque Distribution: (dCq/dX vs. X and J) X J:0.40 0.43 0.46 0.49 0.52 0.55 0.58 0.61 0.64 0.67 0.70 0.30 0.006 0.006 0.006 0.006 0.006 0.006 0.005 0.005 0.005 0.005 0.004 0.45 0.014 0.013 0.013 0.013 0.012 0.012 0.012 0.011 0.011 0.010 0.010 0.60 0.021 0.021 0.021 0.020 0.020 0.020 0.019 0.018 0.017 0.017 0.016 0.70 0.025 0.025 0.025 0.024 0.024 0.024 0.023 0.022 0.021 0.020 0.019 0.027 0.027 0.026 0.026 0.026 0.025 0.024 0.023 0.023 0.022 0.020 0.75 0.80 0.028 0.028 0.028 0.027 0.027 0.026 0.025 0.025 0.024 0.023 0.021 O.85 0.029 0.029 0.029 0.028 0.028 0.027 0.026 0.026 0.025 0.023 0.022 0.90 0.029 0.029 0.029 0.028 0.028 0.027 0.026 0.026 0.025 0.023 0.022 0.95 0.030 0.030 0.030 0.030 0.029 0.029 0.029 0.028 0.028 0.027 0.026 Angles of Attack (Degrees) X J:0.40 0.43 0.46 0.49 0.52 0.55 0.58 0.61 0.64 0.67 0.70

0.30

15.50 14.42 8.60 8.12 6.87 6.14 5.05 4.09 3.35 2.50 1.68

0.45

12.30 7.62 6.75 6.31 5.43 4.65 4.20 3.34 2.66 1.95 1.29

0.60

6.39 6.12 5.35 4.76 4.18 3.73 3.16 2.58 2.07 1.49 0.92

0.70

5.69 5.15 4.64 4.14 3.75 3.25 2.74 2.25 1.77 1.25 0.75

0.75

5.31 4.83 4.37 4.13 3.54 3.04 2.58 2.07 1.64 1.16 0.65

0.80

4.97 4.52 4.33 3.74 3.29 2.84 2.40 1.94 1.51 1.05 0.58

0.85 4.63 4.21 4.05 3.51 3.06 2.63 2.19 1.81 1.38 0.94 0.49

0.90

4.26 4.16 3.64 3.20 2.80 2.39 2.05 1.61 1.23 0.82 0.39

0.95

7.38 6.82 6.25 5.69 5.13 4.58 4.02 3.47 2.91 2.36 1.82

ReynoldsNumber(millions)

X J:0.40 0.43 0.46 0.49 0.52 0.55 0.58 0.61 0.64 0.67 0.70

0.30

0.066 0.064 0.053 0.052 0.050 0.049 0.048 0.047 0.046 0.045 0.044

0.45

0.103 0.082 0.079 0.077 0.075 0.073 0.071 0.069 0.067 0.066 0.064

0.60

0.104 0.103 0.098 0.095 0.093 0.090 0.088 0.086 0.084 0.081 0.079

0.70

0.110 0.106 0.103 0.100 0.097 0.094 0.092 0.089 0.087 0.085 0.082

0.75

0.110 0.106 0.103 0.101 0.097 0.094 0.092 0.089 0.087 0.084 0.082

0.80

0.108 0.105 0.103 0.099 0.096 0.093 0.090 0.088 0.085 0.083 0.081

0.85

0.105 0.102 0.101 0.097 0.093 0.091 0.088 0.086 0.083 0.081 0.079

0.90

0.100 0.099 0.095 0.092 0.089 0.086 0.084 0.082 0.079 0.077 0.075

0.95

0.132 0.123 0.115 0.108 0.102 0.097 0.092 0.088 0.084 0.080 0.077

* Thicknessvalueslimited by availableC1andCd data

for the selected airfoil section.

Appendix E - Manufacturing Plan

Introduction and General Manufacturing Concerns

This appendix details the construction phase of the design project. It presents the assembly breakdown, the major concerns associated with each component, the individual construction responsibilities, the macro schedule for construction, and a brief run through the assembly procedure.

There are a two major concerns that will carry though all phases of the construction process. First, everything must be done to prevent the weight of the aircraft from growing. Through careful planning, the Elite design team has tried to anticipate most of the parts necessary to build the aircraft. Additional pieces will only be added if they are deemed absolutely vital to the success of the technology demonstrator.

Secondly, the center of gravity of the aircraft will be checked throughout the building process. This will help to ensure that the center of gravity ends up in the design location once the construction is complete.

Assembly Breakdown and Major Component Concerns

The assembly of the Elite technology demonstrator is divided into component sections. These components are in turn divided into their substructure components. The breakdown of components is shown below. Also shown in the table are the critical assembly areas or concerns associated with each section.

Fuselage Main wing/fuselage mating Aircraft frame Connection with horizontal stabilator Engine mounts Firewall Servo tray Nose gear mounts Mating of vertical tail Engine nacelle Stiffness of servo tray to avoid Tail cone flexure under control activation Engine vibrations Sturdy nose gear attachment

MainWing

Maintaining airfoil shape

- Wing frame

Mating of two wing halves

- Landinggearmounts

Connection to fuselage

- Ailerons

Sturdy landing gear mounts that resist splitting and tearing of wood

HorizontalStabilator

Maintain airfoil shape

- Stabilatorframe

Solid connection to fuselage that

- Fuselage mounts

still allows for easy control activation

- Control horn placement

Avoid structural flexure during activation Vertical Tail Perpendicular alignment with Tail frame fuselage Rudder Avoid structural flexure during rudder activation Control horn placement Miscellaneous Subsystems Stiffness of control rods Landing gear Smooth and easy control activation Control Rods and Links As one can see for the above table, a major area for concern is the mating of the main aircraft components. In these areas, very. detailed instructions and tolerances must be laid out to insure that all the components will fit once they have been assembled separately. In addition, the different component assembly team must keep in constant communication if any structural modifications are made.

Another concern is the structural response to a control surface deflection. The structure must be carefully constructed to prevent significant structural deformations tinder the loads associated with the surface deflection.

Assembly Teams and Macro Fabrication Schedule

The Elite design team was divided into divisions responsible for the fabrication or the individual major components of the aircraft. Doug Staudmeister and Steve Stem are in charge of the fuselage construction. Jonathan Fay and Dan Avis are heading Lip the wing and stabilator fabrication. Amy Rueter and Tuan Le are building the vertical tail as well as the engine nacelle and tail cone. Below is a macro time schedule for the fabricationprocess.

April 8-10

Cuttingof RawMaterials/Part Construction

Bulkheads

Ribs

SparsandLongerons

Firewall

CompositeBeamConstruction(graphitetapeenforcedbeams discussed in

Section9.0)

April 11-15

Fuselage Construction

Wing Construction(two wings madeconsecutively)

Empennage ConstructionandIntegrationinto Fuselage

April 16-17

MonokotingAircraft

Initial Major Component Mating

April 18

Day to meetpressingand/orunforeseen manufacturing difficulties

April 19

Aircraft Rollout

Brief Manufacturing Assembly Procedure

The airframe assembly begins with the shaping of the top and bottom Fuselage: longerons. Since the Elite incorporates a curved fuselage, the hmgeron_ will have to be pre-soaked in water and then pinned into the desired shape on the full scale plans. These longerons will then have to be allowed to dry, thus forming a permanent curve into the longeron. Once the top longeron has been formed, the bulkheads will be placed along its length except for the rear-most bulkhead. Next the pre-formed bottom longeron will be laid into the bulkheads along with the two side longerons. At this time, the firewall will be added to the airframe and the servo tray can be integrated into the fuselage. With the servo tray complete, the servos can be installed and the control linkages and rods can be laid out inside the aircraft frame. In the area above the wing, a sheet of balsa will extend down along the sides of the fuselage to the wing surface, thus providing a

smoothsaddlematingto thewing. The hardwoodblocksfl)r the main

wing matingcanbeaddedalongwith theengine and nose gear mounts.

This completes the fuselage construction until other major component_ have been completed.

Main Wing: The main wing construction begins with spacing out the ribs in the wing jig one half of the wing at a time. The top, bottom and leading edge spars can be glued into place. The leading edge spar is then sanded into a smooth round shape. The partial wing frame is then removed from wing jig and the trailing edge spar is added to the frame. Next the trailing edge is added to the rear spar at the non-aileron spanwise positions. With half of the wing frame complete, the aileron assembly can be added to the wing with the control wire running along the hinge line. The above steps arc repeated to form the opposite side of the wing. With two completed wing halves, the wing is locked into the desired dihedral angle while the spars from each wing half are cemented together with fiberglass. At this time the front-bottom-center section of the wing is sheeted with balsa to provide a resting place for the batteries. Now the fuselage mating blocks and the main gear blocks can be incorporated into the wing along with the soft balsa rounded wing tips. At this time the wing may be covered. Lastly the main gear are added.

Horizontal Stabilator: The horizontal stabilator frame is constructed just like a miniature wing except that extra care must be taken to build in the desired sweep back angle. Three hardwood blocks are slid into the middle of the graphite rod but not glued to it (the middle hardwood block has the control horn screwed to it). These blocks are the stabilator connection to the fuselage. Hardwood blocks are then glued to the appropriate rib and spar locations in the stabilator frame as the attachment cowlings for the graphite rod. Each half of the stabilator then slides onto the graphite rod and securely glued to it. At this time the stabilator can be covered with monokote. Once the horizontal stabilator has been constructed the outer two hardwood blocks in the middle of the graphite shaft will be glued to the second to last fuselage bulkhead. Now the rear-most bulkhead will be glued to fuselage in effect sandwiching the hardwood blocks between the bulkheads. Then the middle hardwood block is glued to the shaft and the control horn linked to the control rod. Thus the fuselage-horizontal stabilator assembly is complete.

Vertical Tail: The vertical tail and rudder are simple truss structures with rounded leading edges. Once these trusses are complete, they can be covered.

After covering, the two parts can be hinged together. The vertical tail assembly is glued to the top longeron of the fuselage. In addition, the trailing edge spar of the vertical tail extends into the fuselage and connects to the rear bulkhead of the aircraft. The leading edge of the rudder also extends into the fuselage where the control horn is attached.

Engine Nacelle and Tail Cone: These are light-non load carrying coverings that are constructed to complete the sleek fuselage curves at the nose and tail respectively. They are fabricated in the same fashion as the fuselage frame.

Miscellaneous Construction Information: - The glue used in construction is the fast drying cyanoacrylate glue.

Cost Accounting and Control: Each member of the team is responsible for logging his own hours and machine use on the "Tooling Time Sheet" and "Construction Time Sheet." On the following two pages are samples of these two sheets.

CONSTRUCTION TIME NAME DATE #HOURS TASK Page 1

TOOLING TIME SHEET

NAME TOOL TURNON?(Y/N) TIME USED COST

Page2

Appendix F

Flight Validation, Component Test

and

Manufacturing Hours

Flight Validation Testing Review April 21-27, 1994 The Elite Summary: The technology demonstrator was completed and aircraft was noteworthy for the extreme attention to detail in the manufacturing and its attractive "style." It successfully completed its initial taxi tests and it handled very well. The flight tests were rather dramatic.

The first takeoff was successful but after completing a single 180 ° left-hand turn the right wing failed terminating the flight.

Taxi Testin_ Results: Aoril 21. 1994 Ground handling of the aircraft was excellent. Steering was responsive and all the members of the design team were able to control the aircraft. Acceleration at partial throttle was good but no attempts were made to rotate off the nose gear during the taxi tests.

Flight Testin_ Results: Aoril 27. 1994 The final data sheet for the technology demonstrator is attached.

Preflight inspection revealed that the aileron hinges had loosened and the ailerons were reinforced with plastic tape. The nose gear was realigned. For the first flight the aircraft accelerated to takeoff speed, rotated and lifted-off wings level. Rotation to take-off attitude was easily accomplished and the aircraft climbed to approximately 15' and immediately entered left-hand 180" turn. The bank angle appeared to exceed 60 °. Roll response was very good but full aileron and rudder were used in the turn. Immediately upon leaving the turn as the aircraft approached a wings level attitude the right wing (outboard wing in the turn) failed at the root, bent upward approximately 90" and the aircraft crashed.

Post flight inspection of the wing indicated a failure of the upper spar flange (cap) very similar to the component tests. The outboard ribs in bays i and 2 were crushed, the "posts" between the spars in the second bay had separated at their connection points, the web in bay I (the grain in the web was spanwise) appeared to tear along the grain. The failure was localized to the main spar in bays I and 2. The construction team cited the recollection that the balsa used for the upper spar cap was "softer" than that in the component test wing.

Wing Component Static Load Test, April 19,1994 Spring 1994 The Elite Summary: A wing component was tested to failure. The wing was completed (excluding ailerons) and attached to a rigid centerbody in a manner similar to the actual fuselage attachment. The weight of the wing as tested was not provided. The loading was based upon an aircraft weight of 4.84 lb and the wing was designed to a 3g limit load.

This component wing actually was used to conduct two separate tests. An initial load deflection test was conducted in which a @.51b load was applied at each wing rib location from the root to the tip.

This was done for the left wing without monokote and the right wing after it had been monokoted. The bending moment at the root due to the 0.51b load were computed and the deflections at the tip were measured. Both are listed on the attached Table and Plot. During this test audible "cracking" was apparent in the left wing but no visible damage was detected and the wing was not repaired.

During the component test to failure, the left wing (the one already subjected to a static load) failed first with failure of the main spar near the root. Due to the manner in which the wing was mounted to the static test support - a four bolt attachment of the carrythrough to a rigid support- the loading was able to be increased until failure was encountered on theright wing. In both cases the failure mechanism was identical. The top spar failed in what appeared to be a buckling collapse. The most inboard wing bay had a balsa sheet for a spar web but the sheet did not extend from rib to rib. In the next outboard bay three small "posts" were used in lieu of a spar web. It appears as if the rib between the two bays was crushed, the spar separated from the balsa sheet in the first bay (either debonded or splintered) then the unsupported length of the top spar was between the root rib and the first post. Failure occurred in the top spar in this region and the wing failed.

3-g Load Distribution: The approximation to the 3-g ultimate load was applied starting at the root. The load was based upon an assumed aircraft weight of 4.8 lb. The spanwise locations where the loads were applied started 3" from the root and were spaced at 6" intervals. The 3-g load was applied in the increments shown in the table starting at the root.

This processes continued until the left wing failed. The left wing

failed when the total ]oadapplied to both wings was 7.8 ]b. The right

wing failed when the load applied to the right wing only was 5.65 lb.

Spanwise location (distance Load lb

from root

in inches)

3 1.1 {.5, 5, 1}

9 _.1 {.5, 5, _}

_5 1.1 [.5, 5, l}

21 1. {.5, .5 27 .95 {.5, .25,. ,.I) 33 .85 { .5,.25,.1} 39 .75 {.5, .25} 45 .35 {.25,.1} Wing Tip Def]ection: The tip def]ectionwas measured as the ]oadwas increased. The tip deflection is presented for ]eft wing only up to the point where the left wing failed.

Total Load (1 Tip Deflection C_) b)- Both wings 4 , 4.6 1.5 5.6 1.75 6.6 2.25 7.6 3.5 7.8 5.25 Additional Information: Aircraft Weight = 4.84 lb (estimate at this time) Wing Weight = not provided "6 <1) c- O ¢1) ¢1) (1) r.- ._ O0 ,,e,-- c- O cO o °-- ,e..a r- CO -I rr" / t _ -.d

Comparison Between Design and Actual Aircraft Data

Actual Value

Design Value

-- ?

Wing Span -- "

Wing Area ?. c, ." i -- "

Vertical Tail Area 0 ._t_

Horizontal Tail Area 0

Wing Structural Weight (Monokote) , _:

Wing Structural Weight (no Monokote)

) ._%2 .

Fuselage Structural Weight Monokote

Fuselage Structural Weight 'no Monokote)

?,_i: 'b

Vertical Tail Weight (Monokote)

Vertical Tail Weight (no Monokote)

_. i<,,,_, :3 9

Horizontal Tail Weight (Monokote)

.2

Horizontal Tail Weight 'no Monokote)

Landing Gear Weight

Propeller Type

.1

Propeller Weight

Total Aircraft Weight (post-construction)

Total Aircraft Weight (post-flight)

..... if).

CG Location (post-construction)

CG Location (post-flight

471 " ,_ r, d '_ I'?l ' :

Weight of Batteries

• _-- --7'

Please list any other deviations of the technology demonstrater from the original design.

"_7 _CY, xd.,'_ l._ (m_IN_L PA_E m

oF _ _ALn, V

FINAL COST ANALYSIS TYPE OF COST ESTIMATED COST ACTUAL COST FIXED SUBSYSTEMS S462.58 RAW MATERIALS $160.00 ._" 2 r "" ..

PERSONNEL $1800 (180 hrs) TOOUNG $1 50 '_ ' ' WASTE DISPOSAL $ 240 TOTAL COST PER AIRCRAFT $4,410.13 QAI_NAL PA_ Page 1 i : k_

CONSTRUCTION BREAKDOWN

CONSTRUCTION ITEM NO. OF MAN-HOURS % OF TOTAL HOURS WING #1 AND WING #2 57 27.50% MONOKOTE 4 4 21.20 % FUSELAGE 36.75 17.70% STABILATOR 25.75 12.40% MISCELLANEOUS 3 9 19.20% VERTICAL TAIL AND RUDDER 4.25 2.00%

Jolcs

I , # Oi_,ki,M... PA(_IE f_ Page 1