Aircraft Design Push It to the Limit Final

8/12/2019 Aircraft Design Push It to the Limit Final

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Abstract

The Push it to the Limit is an Unmanned Aerial Vehicle (UAV) to be used as a drone for data

acquisition. Namely, the aircraft is fitted with a Fog Aerosol Sampling System (FASS) which,when flown through fog, will record concentration and make-up of the air. The ConceptualDesign will discuss the preliminary design considerations and rough performance estimates.This design uses a puller propeller configuration with a high wing, conventional tail andsquare fuselage. The FASS will be mounted on the underside of the fuselage. Consideringthe mission, maneuverability was the primary concern in developing the conceptual design.


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Who put this thing together? Me, thats who! Who do I trust? Me!

-Antonio Montana

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Contents

0.1 Nomenclature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

1 Conceptual Design 31.1 Proposal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.2 Take-off Weight Estimate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41.3 Wing Loading Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.4 Main Wing Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51.5 Fuselage Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 61.6 Horizontal and Vertical Tail Design . . . . . . . . . . . . . . . . . . . . . . . 71.7 Take-off and Landing Distances . . . . . . . . . . . . . . . . . . . . . . . . . 71.8 Structure Design and Material Selection . . . . . . . . . . . . . . . . . . . . 81.9 Static Stability and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . 101.10 Performance Predicition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

2 Detailed Design 142.1 Wing Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142.2 Fuselage Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

2.3 Empennage Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152.4 Landing Gear and FASS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 152.5 Stability Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

3 Aircraft Fabrication 253.1 Wing Fabrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 253.2 Fuselage Fabrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 263.3 Empanage Fabrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 273.4 Tail Dragger and Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28

4 Flight Testing 29

4.1 Flight Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 294.2 FASS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31

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0.1 Nomenclature

English GreekAR Aspect Ratio Angle of Bank

CDo Drag Polar Coefficient of FrictionCl 2-D Lift CoefficientCL 3-D Lift Coefficient

CLmax Max Lift Coefficientc Chord LengthD Drag ForceDf Fuselage DiameterL Lift ForceLf Fuselage Lengthn Load FactorR Turn Radius

S Wing AreaT Thrust

Tmax Max ThrustV Velocity

Vcruise Cruise VelocityVstall Stall Speed

W WeightW.L. Wing LoadingWTO Take-off Weight

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Chapter 1

Conceptual Design

1.1 Proposal

We propose to design, build and conduct flight tests on a remotely piloted aircraft that col-lects fog concentration data in valleys. The Unmanned Air Vehicle (UAV) will be designedto carry an aerosol sampling system and fly into the steepest valleys possible for data collec-tion. For the design competition, a fog generator will be set up in a field and the UAV willcollect data at two different altitudes. The first will be at approximately 15 feet above thevalley and the second approximately 30 feet. In addition, several configuration constraintswere imposed:

The UAV can be either a tail dragger design with a puller propeller configuration or atricycle landing gear with a pusher propeller configuration.

The engine will be the Kontronik FUN480-33 electric motor with an APC 11x7 E

propeller.

The UAV must have sufficient internal volume to accommodate the radio control re-ceiver, the number of servos needed, the micro-controller board, the flight battery,speed controller and the battery eliminator circuit as well as provide for easy access tothese items.

The Fog Aerosol Sampling System (FASS) device must be securely mounted to thefuselage.

The primary mission of the UAV is to provide a stable platform for the aerosol sampling

system while providing enough maneuverability to safely fly into and out of the valleys.Therefore, the primary design drivers are static and dynamic stability as well as high turnand climb rates. In order to provide a stable platform, a high wing design was chosen sincea high wing aircraft is inherently stabilizing. An airfoil with a high maximum lift coefficientwas chosen so that the plane will be maneuverable enough to fly into and out of the valleys aswell as to turn around within the valleys. In order to collect accurate data, the propeller willbe disengaged briefly during flight to allow clean airflow into the FASS. Therefore, favorableglide characteristics act as a secondary design driver.

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For the fuselage, a basic square cross section was chosen to provide easy mounts for allthe electronic components as well as the landing gear and engine. The fuselage will begin totaper at the wing tip and continue until the tail. Since the aircraft will be a puller-prop UAV,a tail-dragger design will be used along with a conventional tail. The FASS device will bemounted underneath the fuselage since the drag due to the device will provide a stabilizing

moment in pitch. High lift devices will not be necessary since the aircraft has a large thrustto weight ratio and the main wing airfoil has a high lift coefficient. The horizontal andvertical tails will be flat plates to minimize structure weight and manufacturing complexity.

1.2 Take-off Weight Estimate

In order to estimate the take-off weight of the UAV, the following weights in Table 1 wereused for the components of the plane. The total weight assumed 2 small servos for aileronsand 2 large servos for the rudder and elevator. The goal weight for the wing, fuselage, andtail structure will be less than or equal to 1.8 lbf. Past groups have achieved a structure

weight around 1 lbf, however for estimating performance, we used a conservative estimateof the structure weight. During construction, we expect to decrease the final weight of thestructure to around 1 lbf. This change in weight will improve the flight performance fromthe conceptual estimates. The FASS system will be attached with metal clamps that weestimate will weigh 0.05 lbf. The total weight was estimated to be just over 5 lbf which isconsistent with past trends for comparable UAVs.

Table 1.1: Component Weight EstimatesComponent Weight [lbf]

Propeller 0.048

Engine 0.460Engine Mount 0.085

Speed Controller 0.106Landing Gear 0.410

Transmitter Receiver 0.031Large Servo 0.106Small Servo 0.042

Large Battery 0.966Flight Recorder 0.094

GPS Module 0.025FASS 0.750

Metal Clamps 0.050Structure 1.800

Total Weight 5.121

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1.3 Wing Loading Selection

In selecting the wing loading, the primary design drivers for the project were considered.These design drivers are minimum turn radius and a high climb rate in order to maneuverwithin the confines of the proposed valley. Due to the fact that the propeller must be disabled

during data acquisition, the glide capabilities of the aircraft will act as a secondary designdriver.

Fortunately, all three of these design drivers call for low wing loading. For comparableUAVs, low wing loading refers to any wing loading in the range of 18-22 ounces per squarefoot (oz/ft2), or 1.125-1.375 lbf/ft2[2]. Other texts suggest that sport planes, which havefavorable turning and climbing characteristics, have a wing loading between 20-25 oz/ft2

(1.25-1.5625 lbf/ft2)[3].It was decided to aim for a wing loading of approximately 1.2 lbf/ft 2. The following is

the definition of wing loading,

W.L.=W

S (1.1)

Wing loading is a function of gross weight and wing planform area. With an estimatedaircraft weight of approximately 5 lbf, it can be determined that the planform area of thewing must be 4.1667 ft2. As the aircraft will be flown at low subsonic speeds, there is noneed for wing sweep or taper, therefore a rectangular wing planform will be used. Finally,for the purpose of maneuverability, a relatively small aspect ratio of 6 will be used. Withthis final constraint, the dimensions of the wing were determined.

Ultimately, the aspect ratio and wing dimensions were determined based on the desiredwing loading and approximate weight of the aircraft. This will optimize the primary designdrivers of turning radius and climb rate. Although the glide characteristics will suffer due tothe relatively low aspect ratio, this secondary design driver must be compromised in order

to achieve the best possible turning and climbing rates.

1.4 Main Wing Design

Based on the analysis of the optimal wing loading for a UAV, the aspect ratio was chosento be 6 and the wing planform area was calculated to be 4.1667 ft2. The wingspan wascalculated to be 5.0 ft and the chord was found to be 0.8 ft. In order to meet the primarydesign driver of a stable platform for the FASS device, 3 of dihedral will be built into themain wing.

The primary design drivers of the aircraft that affect the airfoil section are high climb

rate and high turn rate. In order to maximize these performance parameters, the airfoilsection chosen was the GOE 498 shown in Figure 1.1. This airfoil was chosen because of itshigh maximum lift coefficient of 1.915 as well as its relatively high stall angle of 14.5 degrees.The airfoil has a thickness of 15.9 % of the chord as well as a 5.5 % camber. The relativelysmall amount of camber will make manufacturing the wing easier since there is not a largeamount of curvature on the bottom.

At low subsonic speeds, the effective Mach number seen by the wing does not need tobe changed and therefore there is no benefit to adding sweep to the wing. The taper ratio

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Figure 1.1: GOE 498 Airfoil Cross-Section

was set to 1.0, which corresponds to a rectangular wing. This configuration will be easierto manufacture and will provide more favorable stall patterns than a highly tapered wing.With these values set, the wingspan, root chord and wing planform could be calculated. Thewingspan was found to be 5.0 ft and tThe chord was found to be 10 in.

With the size of the wing set, the aerodynamic properties of the wing could be foundusing lifting line theory. With the airfoil characteristics and the size of the wing, the wingdrag polar was found to be 0.23. At level flight the lift coefficient was found to be 0.51. Thelift to drag ratio was calculated to be 22.36 and the total drag due to the wing calculated as0.218 lbf.

1.5 Fuselage Design

For the design of the fuselage, the object was to choose a design shape that would be easyto build, hold all the major components (Battery Pack, Servos, DAQ, and GPS) and also

exhibits a low skin drag. One of the biggest factors when determining the fuselage is toreduce the bluff body effects, thus we decided to have three sections to our fuselage: thetapered nosecone covering the engine components, the main fuselage to hold all the majorcomponents, and a rear tapered section providing a moment arm to the tail control surfaces.

Based on measurements of previous senior design aircraft and the width of one and halfbatteries, the height and width of the main section of the fuselage was chosen to be 3.5 inches(in.) x 3.5 in. respectively. The diameter was then taken to be the diagonal distance ofthis fuselage. The total length was chosen based on a fineness ratio (Df/Lf) of 0.11, which

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is typical for many subsonic commercial aircraft as well as other aircraft in previous designyears. Thus, the aircraft length without the nosecone came out to be 45 inches long.

Based on the fact that the wing would be attached above the fuselage, the wing boxwas not considered in determining the volume of the fuselage. The length of the mainfuselage will be based on the main wing positioning. The tapered section of the fuselage

begins immediately after the trailing edge of the wing. Currently the main fuselage lengthis estimated to be 18 inches long, resulting in a square box with a volume of 220 in.3. Aftermeasuring all the major components their combined total volume came out to be 21.8515in.3 (no wires included). This allows for movement of components within the fuselage forfine adjustments to the center of gravity location.

With the previously described geometry, without including the nosecone, the drag overthe fuselage totals roughly 0.021 lbf, which in turn gives an equivalent coefficient of drag forthe fuselage of 0.002192.

1.6 Horizontal and Vertical Tail Design

For ease of design, a conventional tail will be used. This aircraft is functionally comparedto a homebuilt aircraft, so the coefficients of the vertical and horizontal tails (0.04 and 0.5respectively)[1].

In this preliminary stage, the wing is assumed to be mounted at around 11 inches fromthe front of the fuselage, in keeping with past comparable designs. For both the horizontaland vertical tail designs, the surfaces are approximated as flat plates by comparing them tothe very thin NACA 0008 airfoil characteristics.

The vertical tail design has a taper ratio of 0.4 and an aspect ratio of 2, giving a slightlyswept appearance. The horizontal tail has no taper, in keeping with the majority of previousdesigns featuring a rectangular tail, but has a somewhat larger aspect ratio of 2.5.

These design specifications result in a vertical tail area of 0.2688 ft2, and a horizontaltail area of 0.654 ft2 on either side of the fuselage. Taken as a whole, the tail contributes0.016 lbf of drag, which does not raise any problems given the max thrust of around 1.3 lbf.

1.7 Take-off and Landing Distances

The take-off and landing distances for the UAV are estimated from CDo, A, WTO, S, andTmax that were found from previous calculations and research. The one propeller engine setsthe thrust. The wing area remained the same since the flaps dont extend, which leaves thewing area at 4.2 ft2. The rolling friction coefficient for take-off was chosen to best reflectthe surface of the field that the plane will fly on; in this case, firm and dry dirt was chosenwith a=.04. The projected landing gear area was predicted to be around 0.5 by looking atprevious UAVs with the desired landing gear style. A climb angle of 5 degrees was chosenarbitrarily. There is no obstacle to avoid.

From the initial values, the take-off velocity, the dynamic pressure, the wing loading, andthe thrust to weight ratio are calculated. The transition radius as well as the quantities f1andf2 were calculated. From the inputs, it was shown that ground distance was approximately

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109.8 ft, or 70% of the take-off distance. The rotation distance was only 22.7% of the totaltake-off distance, or around 35.6 ft. The transition distance consisted of 14.5% of the take-offdistance while the climb portion did not matter due to a non-existent obstacle height.

Therefore, the total take-off distance came out to be 168.2 ft. This is compared to the144.4 ft originally estimated using the take-off parameter based on historical data. They are

somewhat similar but due to including friction and such the take-off distance became longer.For landing, the approach angle was calculated to be a max of -12.6 degrees. An actual

approach angle of -15 degrees was used. This gave an approach distance of 158.9 ft or 44.4%.The transition distance consisted of 15.7% of the total landing distance, where the free-rolldistance was 28.5% while the breaking distance was 40.7 ft or 11.4%. The thrust used duringlanding was given a reverse thrust of T/W=-1.34. Reverse thrust was considered necessaryto shorten landing distance, although it will not be used in the actual flight test. Thisgave a total landing distance of 572.9 ft with the 1.6 pilot correction factor. If we ignorethe correction factor, we get a more reasonable landing distance of 358.1 ft, which is stillrelatively large for a UAV. Comparing this distance with the previous distance calculatedfrom historical data chosen showed that 494.0 ft was a comparable estimate in relation tothese new calculations.

For this analysis, the maximum lift coefficient used was 1.44. This gave a CLG equalto 0.8Clmax (equal to 1.152). Lift coefficient is very important towards take-off and landingdistances. Looking at the distances gave a general idea of how much space the plane needsto take-off. They however have no impact on the maximum climb rate and turning rate theplane can achieve. The numbers seem a little high, but some error in calculations is expectedespecially since the equations used are not geared towards R/C planes. By watching videosof planes take off and land from previous years shows that these numbers are in fact high.At testing, we will assume a much shorter take-off and landing distance.

1.8 Structure Design and Material Selection

The structural characteristics of the conceptual design reflect an elementary understandingof the loads typically applied to remote controlled airplanes. By observation and estimation,it was determined that the aircraft would be capable of sustaining load factors of between 2and 3. For a 5 lbf aircraft, these load factors imply that the balsa wood wing would have tosustain loads of 10 to 15 lbf.

As structural materials are provided and not chosen, the construction of adequate struc-tural strength is of the utmost importance to the success and performance of the aircraft.Balsa wood and plywood will be used for everything in the structure except for the landing

gear. The fuselage will feature minimal material, with hollow sections to minimize weight.The wing will contain strong spars and sufficient ribs to hold the shape of the airfoil section.For manufacturing purposes, the wing has a rectangular planform with no sweep or taperratio. Using Prandtl.m, a lift distribution over the wing was determined and is shown inFigure 1.2.

In order to properly build dihedral into the wing, it will be built in two sections thatwill be connected in the middle. This junction will be reinforced with fiberglass in order tosustain the large shear forces and bending moments that occur during high loading scenarios

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Figure 1.3: V-n Diagram for Conceptual Design

1.9 Static Stability and Control

The stability of the aircraft is the design component that decides the location of electroniccomponents within the fuselage and wing box. In this preliminary stage it is determinedthat all components will be housed in the forward 18 inches of the fuselage, in the section ofthe structure that does not taper and takes the shape of a 3.5 inch square rectangular prism.

For purposes of control, two larger servo motors will be placed at the very end of thissection and will be connected to pushrods used to control the elevator and the rudder.

The ailerons are controlled by two smaller servo motors mounted in the wing itself for easyattachment and replacement should the need arise.

The engine and propeller are mounted out of the front of the fuselage, and the engineis housed with a separately designed housing piece that attached to the front face of thefuselage box. The landing gear will be mounted at between 2% and 5% of overall fuselagelength. The large servos are attached at the very back of the 18 inch front fuselage section,and the small servos are mounted near the aerodynamic center of the wing structure at about25% of overall length. In this early stage the battery pack is mounted at 10-15%, but theinternal electronic layout will be designed in the detailed stage to allow for as much as 2inches of shifting of the battery pack in either direction as an easy method of adjusted thecenter of gravity location.

Other smaller electronic components are mounted aft of the large battery pack and havelargely insignificant weights. The FASS will be mounted on the planes belly from 5-25% ofoverall length and will be attached with clips to allow for variation in this mounting as asecond method of center of gravity shift.

The fuselage structure is estimated at 0.6 lbf, with a wing structure at 0.9 lbf and acombined tail structure weight of 0.3 lbf. These weights will be greatly refined in the detaildesign stage, but at this point the net weight of the plane, including the FASS, falls at justover the estimated 5 lbs.

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In the current configuration, the plane is longitudinally stable with a static margin of0.032, which is low to allow for maneuverability. Dynamically, a CMA coefficient of 0.2955is also stable. The plane is also directionally stable at a margin of 0.099. To allow for rollstability, a few degrees of dihedral will be implemented into the main wing in the detaileddesign. Currently, it is estimated that (in keeping with past designs) the rudder area will

take up 50% of the vertical tail area and the elevator will be at least 30% of the horizontaltail area.

The majority of past design reports have indicated that upon conclusion of the prelimi-nary design review, the tail surface areas have been increased by as much as 50%. Therefore,in the detailed design, a significant change in tail size and shape is expected.

1.10 Performance Predicition

In order to predict the performance of the aircraft, it was necessary to consider a free bodydiagram of the aircraft in motion. Figure 1.4 below demonstrates the forces acting on an

aircraft engaged in a bank turn.

Figure 1.4: Free Body Diagram of Aircraft in Bank

[4]

By analyzing a force balance, it is shown that for an aircraft flying in steady level flight,engaged in a bank turn, the load factor is a function of the angle of bank according to

n= cos() (1.2)

In the conceptual design, the maximum angle of bank is based on an assumed design loadfactor. Computing the forces in the radial direction and rearranging, it can be shown thatthe radius of the turn is a function of the velocity and the load factor according to

R= V2

g tan()

(1.3)

The minimum radius of turn will occur at the lowest possible velocity with the highestangle of bank, representing the maximum load factor. For a maximum load factor of 3, thesustained turn radius is 9.9 ft, and the plane will be able to make a full 180 turn in under20 ft. Considering a more realistic and conservative load factor of 2, the minimum radiusturn is 16.15 ft, resulting in a full turn with a diameter of 32.30 feet. These predictionsdemonstrate that the aircraft should be capable of flying within the limits of the 90 ft widevalley.

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The climb angle and rate of climb are important performance characteristics to optimize,as the valley walls have a slope of over 60. Unfortunately, the thrust to weight ratio is notlarge enough to achieve a climb angle this high. By considering the excess power of theengine during climb conditions, and balancing the thrust, lift and drag forces acting on theaircraft, the rate of climb and thus climb angle were determined. At a velocity of 44 ft/s,

with an aircraft weight of 5 lbf, thrust of 1.3 lbf and drag of 0.3 lbf, the climb rate wascomputed to be 8.8 ft/s, representing a climb angle of 11.5.

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Chapter 2

Detailed Design

With the conceptual design completed, the next step in the design process is to completea detailed design of the aircraft. In order to improve the perfomance of the aircraft, it wasdecided to add taper to the wings. This reduces the weight of the wings and therefore reduces

the amount of strucure needed to hold the wing in place. The tapered wings also allow forincreased maneuverability and increased rate of climb due to the reduced weight.

2.1 Wing Design

The main wing incorporates a taper ratio of 0.4 for greater maneuverability. The wing has2 degrees of dihedral. It will be constructed out of 16 ribs positioned 4 inches apart, securedtogether by a carbon fiber leading edge spar and two basswood quarter-chord spars. Thinstrips of balsa sheeting will be used for the upper and lower caps on each rib, and a pieceof balsa webbing will connect neighboring ribs. Each rib will have two large lightening holes

that will allow for installation of electronics and pitot tube instrumentation. The leadingedge spar will measure 0.210 inches in diameter, and run through small holes placed tangentto the leading edge radius of the rib. The upper and lower quarter-chord spars will be 1/4inch basswood, and will sit in square cutouts on each rib. Ribs 3 through 8 will have the aft25% of the chord removed. A thin ply spar will be installed along the truncated trailing edgeof these ribs which will allow for ailerons to be mounted with plastic tabs. By nature of thewing taper, the ailerons will also taper toward the wingtips. The ailerons will be constructedfrom 3 inch aileron stock, which will be cut to accommodate the necessary taper.The centerwingbox, measuring 3.5 inches wide atop the fuselage, will be fully sheeted in 1/16 inch balsawood. The remainder of the wing will have a leading edge wrapped in 1/16 balsa wood back

to 30% of the chord length, and a trailing edge wrap will be 25% of the chord length.

2.2 Fuselage Design

The fuselage was designed to minimize the overall weight of the structure and was split intotwo offset halves to increase structural integrity. We started with a base design made of balsawood and then cut out triangle sections to reduce the weight. A simple cut out was made inorder for simplicity in the construction process. Also different sections of the fuselage needed

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to be made out of the light plywood in order to hold heavier parts. These pieces includedthe deck plate below the battery and above the FASS, landing gear, and the tail dragger.Also the firewall behind the engine was made of light plywood.

2.3 Empennage DesignThe tail was designed with 1/4 inch balsa sticks in order to form the plate structure ofthe horizontal and vertical tail surfaces. For the the rudder and elevator we used the balsatrailing edge stock. In order to reduce weight we cut out circular portions of the stock. Sincethe horizontal tail was placed into the rear of the fuselage, the center of the horizontal tailwas made out of balsa stock.

2.4 Landing Gear and FASS

In order to design the landing gear the desired angle of attack of the wing for take off wascalculated based on the weight of the airplane and take off speed. This angle of attack wasbased on 80% of the max lift of the wing. This calculation resulted in the desired heightof the front of the fuselage. The front landing gear was made out of aluminum and thencircular pieces were cut out to reduce the overall weight.

2.5 Stability Considerations

Using Pro/E and the detailed design, the center of gravity of the assembled aircraft was foundto be 10.21 in. from the front of the fuselage. Assuming that the center of lift is at the quarter

chord, the static margin was found to be 0.89 which means the aircraft is longitudinallystable. However, this static margin is not so large as to inhibit maneuverability. As anadded precaution, there is enough room to move the battery up to two inches either forwardor aft to achieve longitudinal static stability.

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Chapter 3

Aircraft Fabrication

The aircraft was constructed using materials purchased from hobby suppliers, chiefly hobbywood (balsa, basswood, and thin plywood), RC Monokote covering, and cyanoacrylate (CA)glue. Other materials for construction, such as servo pushrod connections and fiberglass

tape, were provided to groups in a standard kit. For pieces requiring precise fabrication,design was carried out in PRO/E and actual pieces were cut using a special laser cutter inthe basement of Fitzpatrick Hall.

3.1 Wing Fabrication

Wing fabrication began by determining rib sizing and profile. Using these layouts, 16 ribswere cut (2 each of 8 different sizes to form a 40% taper) with the laser cutter. Theseribs were cut from 18 -inch balsa, with the exception of four ribs that were cut from

18 -inch

plywood. Two plywood ribs on either side of the wing allowed for a secure mounting platform

for 14 -inch-square basswood sticks, into which small aileron servos could be screwed. Using aspecial wing construction jig, the ribs were spaced at a constant distance of 4 inches for theentire wingspan.

With the ribs positioned, 14

-inch-square spar rods were glued to notches cut in the topand bottom of each rib. Since the wing was tapered and the spars were mounted on roughlythe quarter-chord of each rib, the spars did not join perfectly in the wing center. At thispoint, the two wing halves were joined at the spar with a liberal application of fiberglasstape and epoxy. Additionally, the leading edge of the wing held its shape by way of a 14 -inchdiameter hardwood dowel that ran the whole span of the wing.

To reinforce construction, a piece of 14 -inch balsa was added between the spars in the

space between the ribs. This effectively resulted in an I-beam design, which was then furtherstiffened by adding a sandwich piece of 18 -inch plywood glued directly to the spar rods onone side of the I-beam. The leading edge was then wrapped in 1

16-inch balsa sheeting. This

wrap extended to the spars at the quarter-chord.Of the eight ribs, the five outboard ribs on either side had 25% of the aft chord length

removed. Across these ribs, a plywood plate was mounted for the purpose of creating asecure mounting surface for ailerons. The ailerons themselves were created using trailing-edge balsa stock, which was cut and shaped to taper along with the wing. These ailerons

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Figure 3.1: Gold Leader constructing the wing

were mounted to the hardwood surface at four points on either side and were controlled witha single small servo on either side.

To complete the design, the inboard trailing edge was wrapped in 116

-inch balsa, and theentire middle section (between the two most inboard ribs in the center) was sheeted similarlyfrom front to back. A square section in the bottom of the wing center was left open for accessto servo connections and pressure taps leading to a pitot tube, mounted on a wingtip. In

this way, connections to the fuselage could be made quickly, and the wing could be mountedsimply with rubber bands. Following a few further simple additions, such as thin cap stripsadded to the tops and bottoms of ribs and plates added around the servo horns, the wingwas wrapped in Monokote, and the iconic eagle design was added.

3.2 Fuselage Fabrication

The main fuselage structure was designed in PRO/E and cut from large pieces of balsa usingthe laser cutter. This preliminary structure was secured with CA glue. The preliminarystructure consisted of the sides glued to a frontal base portion, followed by attachment of a

back base portion achieved by bending the side pieces and bottom piece to effectively forma seamless taper in the rear section. Support beams were added throughout the fuselagealong the base and in the corners to reinforce the structure to allow for joint strength andthe mounting of a landing gear.

The top portion of the aft section of the fuselage was added after component placementwas determined. This top section included mounting holes for small servo motors thatcontrolled the rudder and elevator. These mounting points were reinforced with small stripsof hardwood for strength. Internally, firewalls were added for torsional rigidity, with a front

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firewall made of hardwood acting as a mounting point for the engine bracket.

Figure 3.2: Laser cutter in operation

The front section of the fuselage was left open at the top, as this was the section overwhich the wing would be mounted. This wing saddle section was reinforced with lightplywood and saddle cushions were added. Short dowels were added on top to act as tiepoints for the rubber bands that would hold the wing in place. A small access door wasdesigned to be placed over components during flight, but this piece was left unattached. The

fuselage was then covered in Monokote, with holes cut out of the Monokote in the aft sectionof the fuselage to allow for airfoil and FASS component access.

3.3 Empanage Fabrication

The stick-built horizontal and vertical tail layouts were determined in PRO/E. For ease ofconstruction, these layouts were plotted full-scale on large sheets of paper to provide anaccurate template for tail construction. The entirety of the tail was constructed using 14 -inch-square balsa sticks, which were cut to specific sizes using the template. The structuresof the horizontal and vertical tails each consisted of nine pieces and were glued together

using the plotted templates for accuracy. The horizontal tail also included a center supportplate cut from 1

4-inch balsa sheet using a band saw.

The rudder and elevator were cut from 14 -inch balsa sheets using a band saw. It wasdetermined that adding lightening holes in these surfaces would carry a high strength penaltyand was not overall beneficial in terms of weight, so the pieces remained solid. The leadingedges of these surfaces were rounded with sandpaper to allow the pieces to hinge. Hinge slotsin these pieces were cut with an X-acto knife for the insertion of CA hinges to connect thesurfaces to the tail structures. Lastly, holes were drilled in the center plate of the horizontal

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tail. This was for the purpose of allowing the bottom sticks of the vertical tail to slot in andthe tail to remain squarely positioned. All parts were then Monokoted.

3.4 Tail Dragger and Gear

A detailed PRO/E drawing of the desired landing gear was produced and presented to theNotre Dame AME machinist, who fabricated the gear structure from aluminum sheeting.Holes for wheel axles were not included, so these were drilled on each end of the gear. Rubberwheels were fixed onto the axles with two set-screw stoppers, and the axles were attached tothe gear structure. Four holes were also drilled into the structures top with a drill press tomount the gear to the fuselage belly.

The tail dragger was constructed with triangular scrap pieces of 18

-inch balsa sheeting.These were glued together with CA glue and then sanded to the desired shape and size. Thepiece was then Monokoted except for the top surface, where it was glued to the fuselage withepoxy to strengthen the attachment against any strong ground forces.

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Chapter 4

Flight Testing

4.1 Flight Data

Flight testing took place over a two day period at the South Bend Radio Control air fieldlocated 15 miles south of Notre Dame. On Day 1 of flight testing, the FASS sensor wasremoved from the aircraft. The first flight served primarily to allow the pilots ample oppor-tunity to become familiar with the aircraft, and to ensure that center of gravity location wascorrect. The second flight tested the planes acrobatic performance, with the pilots perform-ing minimum radius turns, extreme climb and descent rates, and various acrobatic stunts.During Day 2 of flight testing, the FASS sensor was reattached to collect fog data. The firstflight with the FASS attached was used to determine the flight characteristics and acrobaticcapabilities of the aircraft with the new payload. The second and third flights were used tocollect fog data.

According to the Flight Test Program, the pilots gathered flight performance data at

altitudes ranging from 30 to 300 feet. The pilots completed minimum radius turns at lowaltitudes in order to simulate performance in a low level valley. Using data gathered fromDay 2, the minimum instantaneous and sustained turn rates and their corresponding turningradii were calculated. The results are shown below in Table 4.1.

Table 4.1: Predicted and Actual Turning CharacteristicsActual Instantaneous Sustained

Turn Rate 173.5/sec 35.3/secTurn Radius 11.9 feet 96.8 feet

Predicted Instantaneous SustainedTurn Rate 108.2/sec 67.2/secTurn Radius 23.38 feet 37.6 feet

The instantaneous turn rate and radius achieved during the flight test were much betterthan predicted. However, the sustained turn characteristics were not as impressive as thoseexpected according to the spreadsheet.

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Figure 4.1: Completed aircraft prior to testing

Also of importance is the climbing capability of the aircraft. By analyzing several pointsfrom the Day 1 data, a hodograph for climb performance at low level altitude was produced.

Figure 4.2: Hodograph for climb performance

According to Figure 4.2, the maximum rate of climb is 16.81 feet/sec. With a horizontal

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speed of 51.94 feet/sec, the angle of ascent is 17.9. However, to achieve a maximum angleof ascent, the aircraft must fly at vertical and horizontal speeds of 14.17 and 37.4 feet/sec,respectively. This results in a maximum angle of climb of 20.8. By analyzing the descentangles used on final approach for landing, the average descent rate was found to be -10.87feet/sec.

With the data collected, an ideal valley could be created for this plane design to flythrough. Using the wall slope of 2 given, the minimum valley width was found to be 81.8ftusing a smallest turning radius of 96.8ft. By leaving the valley length at 500ft as initiallygiven, the maximum height of the valley was found to be 256.5ft using the maximum sus-tained climb angle of 20.8 starting from 15 ft. Other methods of leaving the valley, likeconstant turning while increasing height and just flying through the middle of the valley andclimbing out, could be used for leaving smaller or taller valleys if necessary. In theory, novalley should be a problem.

Figure 4.3: Hypothetical Valley

4.2 FASS

The Fog Aerosol Sampling Sensor (FASS) was attached for the last two flights. In the firstof the two flights, relative humidity data was collected which indicated ambient relativehumidity at an average of about 59-61%, with spikes as high as 70% and dips as low as 52%.In this flight, the FASS did not indicate that meaningful readings were collected by the fogconcentration sensor. In the second FASS flight, neither sensor detected meaningful changesin ambient humidity and fog concentration. This could have been due to a faulty electrical

connection or issues capturing a usable sample of fog.

4.3 Improvements

Although our plane was successful in its flights, performace and structure could have beenimproved upon. More rudder and elevator surface area could have been added to improvethe aerodynamic performance. Structurally, the fuselage could have been made of more

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plywood, so that less reinforcements were required, thus reducing weight. The ribs had tobe redesigned to use the jig. Construction of the wing could have been more efficient as well.During monokoting, dowel rods to hold the wing to the fuselage should be added after themonokote is done to allow for a flat surface. All in all these minor improvements were notnecessary to produce a spectacular aircraft.

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Bibliography

[1] Corke, Thomas C., Design of Aircraft. Pearson Education, Inc. Upper Saddle River, NJ.Published 2003.

[2] Lennon, A.G. R/C Model Airplane Design. Motorbooks International Publishers &Wholesalers, Inc. Osceola, WI. Published 1986.

[3] Lennon, A.G. R/C Model Aircraft Design: Practical Techniques for Building Better

Models. Air Age Media, Inc. Wilton, CT. Published 1996.

[4] Bower, A.F. Introduction to Dynamics and Vibrations. 3.2 Calculating Forces Re-quired to Cause Prescribed Motion of a Particle. School of Engineering, Brown Univer-sity. Published 2011. http://www.engin.brown.edu/courses/en4/Notes/Particles_PrescribedMotion/Particles_PrescribedMotion.htm

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Aircraft Design Push It to the Limit Final