AIAA DBF '14/'15 Report

School of Engineering

Department of Aerospace Sciences

2014/2015 Design Build FlyFebruary 23 2015

Contents

1 Executive Summary 2

2 Management Summary 32.1 Team Organisation . . . . . . . . . . . . . 32.2 Milestones Chart . . . . . . . . . . . . . . 4

3 Conceptual Design 53.1 Design Constraints . . . . . . . . . . . . . 53.2 Scoring Outline . . . . . . . . . . . . . . . 5

3.2.1 Total Score and Sensitivity . . . . 73.3 Configuration Selection . . . . . . . . . . 8

3.3.1 Fuselage and Boom . . . . . . . . 83.3.2 Empennage . . . . . . . . . . . . . 93.3.3 Wing . . . . . . . . . . . . . . . . . 103.3.4 Propulsion . . . . . . . . . . . . . 103.3.5 Landing Gear . . . . . . . . . . . . 10

4 Preliminary Design 124.1 Design Methodology . . . . . . . . . . . . 12

4.1.1 Airfoil selection . . . . . . . . . . . 124.1.2 Drag estimation: . . . . . . . . . . 134.1.3 Mass estimation: . . . . . . . . . . 15

4.2 Mission constraints . . . . . . . . . . . . . 174.2.1 Takeoff . . . . . . . . . . . . . . . 174.2.2 Maximum speed . . . . . . . . . . 174.2.3 Turn performance . . . . . . . . . 17

4.3 Sizing and tradeoffs . . . . . . . . . . . . 194.4 Estimated aircraft characteristics . . . . . 204.5 Mission performance and Stability analysis 21

4.5.1 Static Stability . . . . . . . . . . . 224.5.2 Tail sizing . . . . . . . . . . . . . . 224.5.3 Dynamic stability . . . . . . . . . . 23

4.6 Propulsion . . . . . . . . . . . . . . . . . . 244.6.1 Methodology . . . . . . . . . . . . 24

4.7 Electronics . . . . . . . . . . . . . . . . . 264.7.1 Battery Selection Factors: . . . . . 26

4.8 Radio System . . . . . . . . . . . . . . . . 27

5 Detail Design 295.1 Aircraft Dimensions and Components . . 295.2 Aerodynamic analysis . . . . . . . . . . . 305.3 Structural characteristics . . . . . . . . . . 33

5.3.1 Load Paths . . . . . . . . . . . . . 335.3.2 Structural Analysis . . . . . . . . . 34

5.4 Aircraft Systems Design, Component Se-lection and Integration . . . . . . . . . . . 36

5.4.1 Fuselage . . . . . . . . . . . . . . 365.4.2 Internal payload bay . . . . . . . . 375.4.3 External ball cage . . . . . . . . . 375.4.4 Wing . . . . . . . . . . . . . . . . . 385.4.5 Wing attachment . . . . . . . . . . 385.4.6 Tail . . . . . . . . . . . . . . . . . . 395.4.7 Tail attachment . . . . . . . . . . . 395.4.8 Fuselage aft section . . . . . . . . 405.4.9 Nose cone . . . . . . . . . . . . . 40

5.5 Payloads Systems Design, ComponentSelection and Integration . . . . . . . . . 415.5.1 Ball drop system . . . . . . . . . . 41

6 Manufacturing Plan and Processes 466.1 Manufacturing Process Selection . . . . . 466.2 Subsystems manufacturing . . . . . . . . 46

6.2.1 Wing manufacturing . . . . . . . . 466.2.2 Fuselage manufacturing . . . . . . 466.2.3 Aft fuselage section manufacturing 466.2.4 Nose cone manufacturing . . . . . 46

6.3 Landing gear manufacturing . . . . . . . . 476.4 Schedule . . . . . . . . . . . . . . . . . . 47

7 Testing Plan 487.1 Testing Objectives . . . . . . . . . . . . . 487.2 Propulsion Test . . . . . . . . . . . . . . . 487.3 Electronics Test . . . . . . . . . . . . . . . 487.4 Structural Testing . . . . . . . . . . . . . . 49

7.4.1 Wing spar and fuselage boomtesting . . . . . . . . . . . . . . . . 49

7.4.2 Landing gear testing . . . . . . . . 497.4.3 Sensor Drop Test . . . . . . . . . . 497.4.4 Flight Test . . . . . . . . . . . . . . 497.4.5 Schedule . . . . . . . . . . . . . . 50

7.5 Flight Checklist . . . . . . . . . . . . . . . 51

8 Performance Results 528.1 Structural performance . . . . . . . . . . . 52

8.1.1 Wing spar and fuselage boomperformance . . . . . . . . . . . . 52

8.1.2 Landing gear performance . . . . 528.2 Electronics Test Results . . . . . . . . . . 528.3 Propulsion Test Results . . . . . . . . . . 538.4 Flight Test Results . . . . . . . . . . . . . 54

UNIVERSITY OF GLASGOW Page 1 of 54

1 Executive Summary

The objective of this report is to describe in detail the design, manufacturing and testing conducted by the Universityof Glasgow AIAA Design, Build and Fly team to create an aircraft that meets the competition requirements. The aimthis year is to create a radio controlled aircraft that can perform the following flight missions: fly as many laps aspossible in four minutes with no payload, carry 5.0 lb (2.3 kg)1 payload for three laps as fast as possible and fly asmany laps as the team desires while dropping a single plastic ball to a specified area each lap. Furthermore, in theground mission, the team will have to load the specific payload which is three pieces of lumber into the aircraft onthe ground, secure the payload and remove it. Then load the amount of plastic balls intended to be carried in thefinal mission into the aircraft while securing them properly. The ground mission has to be completed as quickly aspossible.

The total score for the team is obtained from three components: Total Mission Score, Written Report Score andRated Aircraft Cost (RAC). Total mission score combines flight and ground mission scores. RAC includes the aircraftempty weight and the number of servos installed. It is also set that the aircraft has to be able to takeoff and landwithin 60 ft (18m) and the battery pack cannot weigh more than 2.0 lb (0.9 kg).

From these requirements it was determined through competition score analysis that empty weight and number ofservos were the two driving factors in the design, followed by the ability to load and unload the aircraft easily, carry alarge number of plastic balls and to have high top speed. For example, if the team goes from having four servos tofive servos, the total mission score will go down by 20%. This is a deduction that is virtually impossible to compensatewith other improvements. Therefore, the team decided to fly the aircraft without a rudder and to only have a singleengine, since ESC would also be counted as a servo.

It was evaluated that the team could not carry a number plastic balls such that their total mass would be higher thanthe mass of the three pieces of lumber. Hence, flight mission 2 was found to be the most demanding in terms oftakeoff and landing requirements due to the highest mass payload required. This dictated the aircraft motor’s powerand wing area combination requirements, consequently overpowering it for flight missions 1 and 3. This, however,enabled the team to choose higher pitch propeller to utilize the excess power and increase the top speed for flightmission 1. It also also allowed to operate the motor at maximum efficiency level for mission 3 extending the range andthus increasing the number of laps that can be flown. Greater wing area also made sharper turns possible decreasingthe lap time.

The team’s final design was a 1.45 kg empty weight midwing monoplane with maximum flight speed of 23m s−1. Themain construction materials were carbon fiber and foam. It can carry six plastic balls, takeoff within 18m carryingmission 2 payload of 2.3 kg. The aircraft is able to complete 6 laps in four minutes for mission 1 and can fly 3 laps formission 2 in 2 minutes.

1Imperial Units will be quoted in this report when some competition specification is also quoted in those units. Otherwise metric system isused.


2 Management Summary

2.1 Team Organisation

University of Glasgow AIAA Design, Build and Fly team has members from first to fourth year undergraduates.Overall, the team has 21 members from 9 countries, supervised by faculty advisor. All work was done by the studentsand on top of regular university curriculum. This was the first time when University of Glasgow has entered thiscompetition and consequently the team management had to start from acquiring equipment, tools, workshop andwork their way up step by step. The organisational structure of the team is presented in Figure 1.

Arturs Jasjukevics 4Project Manager

Aerodynamics

Julius Bartasevicius 4Sifoyiannos Nevradakis 2Pavlina Dimitrova 3 Gani Petelov 4

Aki Ruohonen 4Technical Lead

Deep Pandya 5Business Manager

Vladislav Andrijako 4Treasurer

Structures

Kiril Boychev 3Blaga Todorova 3Junaid Ashraf 1Stoyan Barbukov 2Christopher Logie 3

Electronics

David Thrulbeck 4Seifallah ElTayeb 2Thomas Timmons 1Benjamin Gregg-Smith 1

Propulsion

Angel Zarev 4Euan McLean 4Dragomir Kamov 3

Julius Bartasevicius 4Chief Designer

Dr Hossein Zare-BehtashFaculty Advisor

Thomas Timmons 1Assistant Treasurer

Figure 1: Organizational Chart of the team. Number after each name indicated year of study

Team meetings were scheduled every week and utilized online forum and Google Drive as main lines of communi-cation and information sharing. Project manager was responsible for setting up the master schedule, overall projectgoals and took the final responsibility for all major decisions. Technical lead assisted project manager, was a primarysupervisor for the project’s technical side and managed project workshop, tools, equipment etc. The chief designerreported to the technical lead and also to the project manager to provide an extensive checking system for all projectcritical decisions.

Finance team worked mostly independently from the rest of the team and took the responsibility of financing theproject and maintaining public relations. This was especially important since University of Glasgow participates inthe competition for the first time and the team had to establish their name and make themselves visible for potentialsponsors from scratch.

Furthermore, the team was divided into four sub-teams: Aerodynamics, Propulsion, Electronics and Structures. Eachof those had a designated sub-team leader, who would manage the activities of the sub-team and report to the ChiefEngineer and through him to Project Manager and Technical Lead. Outline of sub-team tasks is as follows:

• Aerodynamics: Sizing of aircraft components, performance and stability estimates• Propulsion: Motor, ESC and Propeller sizing, testing and selection• Electronics: Aircraft battery, servos, receiver and transmitter sizing, testing and selection


• Structures: Aircraft fuselage, landing gear, empennage and drop mechanism construction and testing

2.2 Milestones Chart

Project manager created a general schedule for the project that identified all major steps needed to produce a finaldesign. This included everything from early design to flight testing. Milestones chart is shown in Figure 1.

Activity

A/C Design Planned Concept design Actual

Preliminary design Due date /

Prelim. review Current date

Detailed design

Design freeze

Manufacturing

Prototype 1

Prototype 2

Final Design

Testing

Propulsion test

Component test

Flight test Prototype 1

Flight test Prototype 2

Flight test Final Design

Report

Draft

Edit

Due date

AIAA DBF competition

LegendMarch AprilOctober November December January February

Winter

Exam

Session

Table 1: Master schedule of the project.

Due to the fact that this is the first year the University of Glasgow participates in DBF, the planned and actual scheduledisagrees on some occasions. Most delays were caused by administrative issues, such as late receipt of funding,issues ordering tools and equipment, and inability to use certain facilities due to long queuing times. For next year,the team has been accumulating ”lessons learned” style information to preserve experience that will be used inforthcoming competitions.


3 Conceptual Design

In the conceptual design phase the mission rules and requirements were analyzed to determine the major designfeatures of the aircraft that would maximize the team score. Sensitivity studies were carried out to quantify thecontribution of each section of the competition and to find an optimal compromise for the aircraft configuration.

3.1 Design Constraints

The competition rules set several direct limitations on the possible design of the aircraft. Most important of theseincluded:

1. Maximum battery mass is 2 lb (0.9 kg) of NiCd or NiMH type2. Runway length for takeoff and landing is 60 ft (18 m)3. Mission has to be conducted without significant damage to the aircraft4. Complete all the missions:

(a) Load and unload payloads of mission 2 and 3 in less than 5 min to avoid heavy penalty(b) Be able to carry flight mission 2 payload (5 lb (2.3 kg)) over 3 laps internally(c) Carry at least one plastic ball externally for mission 3 that can be dropped on command

5. Aircraft has to withstand lifting from the wingtips when fully loaded6. Payloads must be secured well and ball drop mechanism designed such that the center of gravity (CG) stays

within limits7. Intentional aerobatic maneuvers are not allowed

The competition course layout with runway, turns and drop zone is shown in Figure 2. Total length of the course isapproximately 3000 ft (1 km).

Figure 2: The flight course layout

3.2 Scoring Outline

Total competition score is given by

Score =WRS · TMS

RAC


where WRS is Written Report Score, TMS is Total Mission Score and RAC is Rated Aircraft Cost. Written ReportScore is based on the quality of the technical report submitted by the team in February.

Rated Aircraft Cost RAC is found from

RAC = EW ·NServo

where Empty weight, EW = Max(M1,M2,M3) hence the highest empty weight for any mission configuration andNServo is the number of servos installed. Definition of a servo for the competition includes, but is not limited to

• Conventional radio controlled servo actuator• Speed controller• Electric motor not used for propulsion• Solenoid actuator• Electric relay

Total Mission Score Total Mission Score is defined as

TMS = GS · FS

where GS is Ground Score and FS is Flight Score

Ground Score Ground Score is determined by the time taken to perform the ground mission. In the mission theteam has to load flight mission 2 payload into the aircraft and secure it. Then the timer will be paused and the payloadsecuring is checked. After this, timer will be resumed while the team loads the number of plastic balls declared in thetechnical inspection for flight mission 3 into the aircraft and secures them. The time is stopped and final check of theloading is done. The score is then calculated from

GS =Fastest Loading Time

Loading TimeUG

where Loading TimeUG is time achieved by the University of Glasgow.

Flight Score Flight score can be calculated from

FS = M1 +M2 +M3

where M1 to M3 denote individual flight mission scores

Mission 1 Ferry Flight Fly 4 min around the course without a payload to complete as many laps as possible. Scoreis given by

M1 = 2 ·Number of Laps FlownUGM1

Number of Flaps FlownM1

Mission 2 Sensor Package Transport Mission Fly three laps around the course with 4.5 ′′ x 5.5 ′′ x 10 ′′ (11 cm x14 cm x 25 cm) stack of pine boards weighing approximately 5.0 lb (2.3 kg) as fast as possible. Number of points isthen

M2 = 4 · Fastest TimeTimeUG


Mission 3 Sensor Drop Mission Fly as many laps as desired by the team while dropping one plastic ball of 3.759 ′′

(95.48mm) in diameter, weighing 1 oz (30 g) to a designated area from the aircraft each lap. Mission score thenbecomes

M3 = 6 ·Number of Laps FlownUGM3


3.2.1 Total Score and Sensitivity

When the terms in the total competition score are now expanded it can be observed that

Score =WRS · TMS

RAC=

=WRS · Fastest Loading Time

Loading TimeUG·(2 · Number of Laps FlownUGM1

Number of Flaps FlownM1+ 4 · Fastest Time

TimeUG+ 6 · Number of Laps FlownUGM3


)EW ·NServo

Therefore the variables for sensitivity analysis are the ground and flight missions performance ratios Fastest TimeTimeUG

etc.,empty weight and number of servos as those are the parameters that are determined by aircraft performance Writtenreport score is not affected by aircraft performance and thus will not be considered in sensitivity analysis. Sensitivityplot is shown in Figure 3. It now becomes clear that reduction in empty mass and in the number of servos have the

-40 -30 -20 -10 0 10 20 30 40-40

-20

0

20

40

60

80

Parameter change, %

Tot

al s

core

cha

nge,

%

Empty massMission 1 scoreMission 2 scoreMission 3 scoreNumber of servosGround score

Figure 3: Sensitivity Plot of the mission requirements

largest effect to the total competition score since their product divides the entire competition score expression. Forinstance 10% increase in empty weight could only be compensated by about 20% increase in mission 3 score or


30% increase in mission 2 score. Furthermore ground score plays a vital part in the total score since it multiplies allthe flight mission scores. Thus if the team scores 0.2 in ground mission, 80% of the flight mission score will be lostcompared to the fastest team. It should be noted that adding a servo causes very sharp drop in the total competitionscore as the percentage change will be very high when moving from, say, 4 to 5 servos.

Last, the aircraft needs to be designed to perform all missions, thus 60 ft takeoff distance and 3 lap range have to bemet at maximum payload, i.e. M2 payload. The requirements are translated to aircraft requirements in Table 2.

Importance Mission Requirement Aircraft Requirement

1 60 ft Takeoff and Landing with M2 pay-load

High lift and Thrust-to-Weight

2 Minimize EW Use of competitive materials (CarbonFiber etc.), optimized design

3 Minimize NServo Minimize number of ESC, systemsthat require servos(flight control, sensordrop), motors not used for propulsion etc.

4 Fast M2 & M3 loading and unloading se-curely

Easy access to internal payload bay,Simple ball loading mechanism, Straight-forward securing systems.

Table 2: Mission requirements translated to aircraft requirements

3.3 Configuration Selection

One of the main design challenges was to design a payload drop mechanism that would not require an extra servo.Initially it was thought to use a device that would drop a ball based on dynamic pressure. This, however, had to beabandoned because of reliability concerns the changing wind conditions might bring.

The following idea was to perform an aileron roll that would cause a ball to fall. This idea dictated that the balls wouldneed to be dropped from the top of the aircraft and that tolerable rolling characteristics could be achieved. Intentionalaerobatics maneuvers were, however, forbidden after rules update and the team had to drop this idea as well. It wasdecided then to have a more conservative approach and have a dedicated servo to operate the drop system.

3.3.1 Fuselage and Boom

M2 Payload

M3 Payload

Flight direction

M3 Payload

drop direction

Figure 4: Fuselage and Boom general configuration


Configuration selection started from the most fundamental aspects of aircraft layout. Since empty weight has a veryhigh importance in scoring, a minimalistic approach was taken for the fuselage design. This meant that the fuselageneeded to have just enough space for the flight mission 2 payload carried internally and for up to 8 plastic balls(which was defined from endurance limitations) for flight mission 3 to be carried externally. Furthermore the plannedusage of composite materials as much as possible shaped the aircraft such that flat composite plates could be usedas much as possible. Film covering would be used to minimize time consuming molding, while at the same timemaintaining acceptable aerodynamics characteristics.

Those requirements led to a conceptual design with a carbon main boom to which all aircraft components were fixed.Ball cage that holds the plastic balls was positioned on the bottom side of the boom and the balls would be droppedfrom the back of the cage. This was done to avoid collisions with the fuselage, empennage and propeller. The cargocompartment for mission 2 payload was mounted above the boom to avoid excessive length and to limit center ofgravity movement in mission 3. This configuration is presented in Figure 4.

The conceptual thought that concerned the fuselage layout was aiming for a design that would be as aerodynamicallyeffective as possible. Effectiveness in this case meant the reduction of drag caused by the fuselage, its contributionto the lift coefficient and its optimum functionality in each mission. These requirements by themselves led to a designbased on an aircraft’s wing airfoil. Based on experimental data and wind tunnel tests on different types of airfoils andfuselages, and while observing actual designs of transport aircraft at the same time, the fuselage was designed in away that resembled a semi symmetrical airfoil. Once an initial approach was architectured and pictured, details thatwould play an important role to its functionality had to be added, with the soothing of the surface from sharp edgesbeing the most important.

3.3.2 Empennage

(a) T-tail (b) V-tail (c) Conventional tail

Figure 5: Different tail configurations

The driving requirements for empennage design were lightweight, simplicity of manufacture and effective provision ofcontrol. The types of tails considered were conventional, V- and T-tails as seen in Figure 5. T-tail offers good clearancefrom wing interference by placing the horizontal tail high above the wing thus enabling smaller area. This, however,forces vertical stabiliser to be very rugged so that it can take the loads created by the horizontal tail positioned on topof it. Also, at high angles of attack, tail can be shadowed by the fuselage causing deep stall that can be extremelydifficult to recover from.

V-tail can reduce drag of the aircraft by reducing the amount of intersections that cause aerodynamic interferenceas compared to conventional or T-tail. However V-tail also creates strong coupling of lateral and longitudinal control.This coupling can create flight behaviour that is difficult to handle when the aircraft does not have a rudder or in highgust situations. Lack of rudder control means that V-tail would lose one of its advantages of enabling pitch and yawcontrol with high aerodynamic efficiency. Hence it is ineffective and complex solution if only pitch control is required.

Conventional tail is similar to T-tail but a horizontal stabiliser is connected to the fuselage/boom rather than positionedon top of the vertical stabilizer. When the horizontal tail is close to the plane of the wing, downwash of the wing willsignificantly reduce efficiency of the tail increasing required tail area as compared to an ideal situation. Advantage isthat the tail can be connected directly to the fuselage/boom which is a simple solution and does not require strong


vertical tail to hold the horizontal tail. Thus conventional tail was found to be the best compromise.

3.3.3 Wing

(a) High wing (b) Low Wing (c) Mid Wing

Figure 6: Different wing configurations

The position of the wing was largely dictated by the location of the main boom. The positions considered are illustratedin Figure 6. The boom is the main load carrying structure of the aircraft and it would be inefficient to attach the wing toanything else since this would require heavy strengthening of the parts that would carry the wing. The boom location,on the other hand, was dictated by the requirement to carry plastic ball externally and timber internally in separatecompartments and at the same time limit length, height and CG shift to minimum.

Mounting the wing low would have enabled the team to take better advantage of ground effect which is very importantdue to strict takeoff and landing distance limitations. However, the reasons stated earlier outweighed the advantagesthat could be obtained from a low wing configuration. Mass especially would have become an issue if the winghad been attached, for example, to the bottom of the ball cage. High wing could have provided extra stability byplacing the wing above the CG, however high wing configuration would have required strengthening of the top cargocompartment. Due to these reasons a mid-wing configuration was selected.

3.3.4 Propulsion

Selection of propulsion configuration was made relatively straight forward by the combination of mission 3 sensordrop requirement and the servo count cost which made every ESC to count as a servo. Thus minimal number ofESCs and a good clearance for dropped balls were the two main factors. Two motors could be powered with a singleESC but this would not provide any significant advantages such as differential thrust control and the mass of thesystem would be most likely higher than with a single motor. Furthermore pusher configuration was quickly discardeddue to the fact that a ball could hit a pusher propeller when ejected backwards and propeller ground clearance issueson takeoff and landing. Therefore single motor tractor configuration was chosen.

3.3.5 Landing Gear

(a) Tricycle (b) Tail Dragger (c) Bicycle

Figure 7: Different Landing Gear configurations

The main requirements for landing gear design were acceptable ground clearance, light weight, good stability &


handling and low drag. Evaluated configurations are presented in Figure 7. Tail dragger configuration is generallythe lightest option since the tail gear can be very light due to the small load on it. However ground clearance was anissue because the low position and long length of the ball cage. This would make the main landing gear very tall andcause too high ground angle of attack which in windy conditions causes significant problems. Also the aircraft CGis aft of the main landing gear which makes this arrangement inherently unstable thus the pilot has to use rudder tobalance the aircraft.

Bicycle gear would require attaching the wheels to the ball cage which necessitated more mass to strengthen it tocarry the weight. Also for ground stability, small wheels would be required at each wingtip, leading to added drag andmass. This solution could be useful when extremely high aspect ratio is used to support the wing upon landing. Thisdesign requires very long runways, since the aircraft can not be rolled due to the aft wheel being far behind the CG.

Tricycle gear has moderate weight but provides good ground handling qualities. Ground angle of attack can alsobe set to a more manageable level for high winds than with a tail dragger configuration. The inherent stability of atricycle gear allows relatively high sideslip angle for landing and thus permits tolerable landing characteristics for arudderless aircraft. Hence tricycle gear was selected.


4 Preliminary Design

4.1 Design Methodology

The constraint analysis method was used to design the preliminary configuration of the aircraft. Flying mission 2 wasdecided to be the most demanding one. This is because it has the biggest payload and also the aircraft has to flyat maximum speed. Mission 3 needs more range, but in terms of power for takeoff, it does not require as much asmission 2. Hence the baseline configuration was designed for mission 2.

The method employs analytical equations of power loading (P/W) as a function of wing loading (W/S) for differentmission segments. The mission segment performance is influenced by various constraints, in this case being eithercompetition requirements, environmental conditions or reasonable estimates.

4.1.1 Airfoil selection

With the two main design goals set, the number of servos to be minimum and also mass to be minimum, the airfoilselection was constrained. It was decided not to use any high lift devices, hence a big maximum lift coefficient wasthe number one criteria for the airfoil selection. Also, it appeared to be one of the biggest influences on minimummass for takeoff, hence a coefficient in the range of 1.5-1.7 was the first constraint used. Bigger lift coefficients wouldresult in very high trim drag for missions 1 and 3 since the takeoff weight is much smaller than in mission 2.

The second selection criterion was for the airfoil to be optimized for low Reynolds number flow, usually being lessthan 500,000. Such flow aerodynamics, when it is mostly laminar, require great attention for a couple of reasons.First, hysteresis is often observed on airfoils operating in such regime. Hysteresis, flow’s dependency on its previoushistory, is particularly important for stall recoveries, spin flight or high gust conditions. This unsteady phenomenonhighly increases the drag and decreases the lift when the angle of attack of an airfoil is reduced after stalling, pos-sibly resulting in loss of the aircraft. Hysteresis is strongly coupled with the second phenomenon of low Reynoldsnumber airfoils - laminar separation bubbles. Laminar separation bubbles form when airflow that is laminar separatesbecause of high adverse pressure gradients and when separated, transitions to turbulent flow. It then curves backand reattaches to the surface, creating a shallow region of reversed flow. Such bubbles increase drag, which canbecome a few magnitudes bigger than the drag of the airfoil without the bubble, and are to be avoided.

The phenomena of laminar separation bubbles can be controlled with selecting the right transition location during theairfoil design phase with transition ramps or with externally fitted turbulators.2 After analysing the list of low Reynoldsnumber airfoils, three options were chosen. Selig SG6043, which was designed for small wind turbines, WortmannFX63-137, which originally was designed for human powered airplane use but also adapted by airplane modellers,and Douglas LA203A, another high lift airfoil. The selected airfoils are shown in Figure 8. The 2D polars can be foundin Figure 9. For cruise lift coefficient ranging between 0.11 to 0.42 for various missions, the Wortmann FX63-137had the lowest drag from the three airfoils. Even though it has a very thin trailing edge that might cause problems formanufacturing, it was decided that with careful use of carbon fiber ribs it could be avoided. Hence the selection wasconcluded with the Wortmann FX63-137 as the airfoil.

2Michael S. Selig. “Low Reynolds Number Airfoil Design Lecture Notes”. In: the von Karman Institute for Fluid Dynamics, Lecture Series.2003.


Figure 8: Three airfoil profiles considered in the final stage. Red is Wortmann FX63-137, blue is Selig SG6043 and green isDouglas LA203A

(a) ClCd

vs Angle of Attack (b) Cl vs Cd

(c) Cm vs Angle of Attack (d) Cl vs Angle of Attack

Figure 9: Aerodynamic graphs of the final three airfoils.

4.1.2 Drag estimation:

Zero-lift drag buildup model from Roskam3 was used to determine drag component by component using 14ms−1,assumed average flight speed, as a reference.

3J. Roskam. Airplane Design: Part VI: Layout Design of Cockpit, Fuselage, Wing and Empennage : Cutaways and Inboard Profiles. AirplaneDesign. DARcorporation.


First, wetted areas of the exposed components were calculated, approximating the fuselage as a box a bit biggerthan the mission two payload. Then skin friction coefficients were estimated, depending on the component locationand the Reynolds number of the flow over the component. Only the tail was assumed to have a laminar boundarylayer, since the fuselage is in the wake of the propeller and the wing was also partially covered by the wake.

Form factor, which defines the interference and pressure drag was then included for each component. For the wingand tail it mainly depends on its thickness ratio and for fuselage it depends on its fineness ratio. For the landinggear the drag coefficients suggested by Hoerner4 were used. All the equivalent parasite drag areas and zero-lift dragcoefficients are tabulated in Table 3.

Component Wetted area, m2 Form Factor Equivalent para-site drag area,m2

Parasite drag co-efficient

Wing 0.7801 1.3012 0.0061 0.0143Tail (horizontal and ver-tical)

0.1659 1.1 0.0015 0.0034

Fuselage 0.2985 4.0594 0.0065 0.0153Landing gear - - 0.0016 0.0036Total - - 0.0157 0.0367Total, including a factorof 1.06

- - 0.0166 0.0389

Table 3: Aircraft Drag Buildup

Using XFLR5, average Oswald’s efficiency factor (e = 0.949) for the wing was estimated. Lift-induced drag then wascalculated and a drag polar for complete aircraft was created. It can be found in figure 10.

4S.F. Hoerner. Fluid-dynamic drag: practical information on aerodynamic drag and hydrodynamic resistance. Hoerner Fluid Dynamics, 1965.


Figure 10: Drag Polar of the complete aircraft

4.1.3 Mass estimation:

List of aircraft parts and components was created and mass of each was estimated. Fuselage mass was based onthe previous reports which had similar conceptual design. Also, 10 g were added for every ball the aircraft is going tocarry, because the ball cage gets heavier. Three component masses were left as functions of different parameters:wing, motor and battery.

Wing density per area was first estimated using previous reports of similar configuration and aspect ratio wings, andlater refined after a CAD model was finalized. However, the first estimate proved to be quite close to the actual value.

Some 300 motors were put into a database and a graph of motor mass versus power was created, as shown in Figure11. Especially in the lower power range, the graph is linear with the deviation increasing as the power increases. Afterfiltering out the overpowered motors according to estimations, the motor mass unknown was substituted by a linearfunction with required power as a parameter.


Figure 11: Motor mass to Power graph

Finally, battery mass was dependent on the required power output and the required capacity. However, the capacityrequirement for mission 3, where battery endurance is required, was a smaller constraint than the power requirementin mission 2. After analyzing a few different cells that would fit the performance range, a mathematical model wasestablished, resulting in a battery mass function with battery’s C rating, maximum power required and the energydensity of the battery chemical (80Wh kg−1) as input parameters.

Battery mass =Maximum power required

C rating · Chemical energy density

The preliminary component masses are tabulated in table 4. With these estimated, mass could be calculated forevery point on the power requirement graph.

Component Mass, kg

Servo 0.03Receiver + receiver battery + Fuse 0.06ESC 0.05Propeller 0.02Boom 0.12Fuselage 0.150 + Nballs· 0.01Landing gear 0.11Empennage 0.11Motor M2 mass · PW / 3534.6Wing M2 mass · 0.73945 / WSBattery Maximum power required / 800

Table 4: Initial Mass build up estimations


4.2 Mission constraints

The 3 most influential mission segments were considered- takeoff, maximum speed and turn performance. Othertwo, cruise for maximum range and climb were found to be less demanding.

4.2.1 Takeoff

Takeoff performance initially was the most influential segment. It dictates the minimum wing size with selectedmaximum lift coefficient. The bigger the Clmax , the lower the stall speed and wing area, decreasing the total mass.

Takeoff also greatly depends on the available field length. A safety factor of 1.25 was used to make sure that it ispossible to take off in the given field of 18.3m.

4.2.2 Maximum speed

The aircraft is supposed to fly at maximum speed during the timed missions. However, since mission 2 score is lessimportant than the empty mass of the aircraft, it was decided not to chase a high maximum speed and only makesure that in windy conditions the aircraft is able to move forward. Research on the environment of the competitionlocation was done for maximum wind speeds. Last year, the winds in Tucson, AZ reached 240 kmh−1. However, flyingagainst such wind for an RC airplane is impossible, so only the wind speeds of three years before were considered,averaging at 70 kmh−1. As a consequence, the maximum achievable speed was chosen to be 82.8 kmh−1 (23m s−1).

4.2.3 Turn performance

It was first decided that a high aspect ratio wing is going to be used for smaller induced drag and better cruiseperformance. This requires higher structural stiffness, especially for the wing root, since bending moments areincreased. Most of the previous reports suggested maximum turn load of 5-6 G’s. It was estimated that a load of5.5G will be achieved during the maximum G turn. However, the actual loads can only be estimated after test flights.

The mission requirements are summarized in a Table 5 and added safety factors can be found in Figure 6. The codeoperating principle is shown in Figure 12.

Design constraints Parameters for code input:

TakeoffFIELD LENGTH Cl max = 1.63 Estimated maximum coefficient of lift (WORTMANN FX 63-137)MAXIMUM ClMAX

roll fric = 0.025 Ground roll friction coefficientt grun = 5 Ground run time, s

Maximum SpeedWIND SPEED cruise h = 15 Desired cruise altitude, m

V max = 23 Estimated maximum speed, ms−1

TurnsTURN SPEED V turn1 = 14 Turn velocity, ms−1

GeneralVARIOUS AR = 8 Estimated aspect ratio

N balls = 6 Number of balls, an initial guessh 0 = 680 Ground altitude of Tucson, sT 0 = 20 Average temperature in Tucson, ◦C

Table 5: Mission requirements


Geometry inputs:Taper, AR, Volume

coefficients, Tail ARs

Aerodynamic inputs: CL_max

Design requirements:

Takeoff distance, Max speed, Turn

speed

Payload inputs:Number of balls

carried, mission 2 payload mass

Typical component

masses

Safety factors

Preliminary aircraft mass

Preliminary drag buildup

Initialize wing loading and

power loading matrix

Wing, motor, and battery densities

Adjust model mass for different

power loadings and wing loadings

Adjust drag for every matrix

point

Takeoff Performance

Turn Performance

Max Speed Performance

Calculate energy required

for all the missions

Battery mass required the same as predicted?

No

Yes

Empty mass matrix

compiled

Minimum mass configuration

Figure 12: Operating principle of the optimization code. Blue represents inputs, grey- outputs. Red shows the iteration loop of thecode.

Maximum cruise range performance was also considered, but it didn’t imply any requirements higher than the missionsegments mentioned above. Climb angle was chosen so that it does not change the power requirement too, resultingin 8◦, which seemed to be sufficient. The resulting power loading and wing loading (also wing cubic loading) curveswere found and a model and calculated empty masses were incorporated into the graph. The resulting graph can befound in Figure 13.

Safety Factors

S takeoff = 60 / 1.25 Takeoff distance (ft), 1.25 divider as a safety factorbattery SF = 1.2 Battery mass multiplier for safetyeta prop max = 0.7 Estimated propeller efficiencyeta moto max = 0.6 Estimated motor efficiencyT hot = 30 Hot day temperature at Tucson, AZ, ◦C

Table 6: Safety Factors of the optimization code


4 6 8 10 12

20

40

60

80

100

120

140

160

180

1.5

1.5

2

2

222.

52.

52.

5

33

3

Empty Mass versus Power Loading versus Wing Loading

Wing Loading, kg/m2

Pow

er L

oadi

ng, W

/kg

4 6 8 10 12 14 16 18 20 22 24

20

40

60

80

100

120

140

160

180

1.51.5

2

2

2

2.5

2.5

2.5

3

3

3

Empty Mass versus Power Loading versus Wing Cubic Loading

Wing Cubic Loading, kg/m3

Pow

er L

oadi

ng, W

/kg

Empty mass, kgTakeoffClimb Performance at 8 degreesTurn speed, 14m/sMax speed, 23m/sMax Range PerformanceMinimum EM Point

Figure 13: Power Loading vs Wing Loading and Power Loading vs Wing Cubic Loading. Note that the mass values below theturn constraint line have no physical significance. This is due to optimization code estimating that the airplane stalls when poweravailable is less than that needed for a turn. However, this does not influence the minimum mass point in the design space in anyway. Also the mass lines are discrete, because of battery mass increasing not linearly, but as an integer multiplier of battery cell

mass.

4.3 Sizing and tradeoffs

As can be seen in Figure 13, lowest empty mass point was automatically marked and all the geometry and aerody-namic characteristics for that point were generated. However, the suggested values for wing loading were discussedin a bit more detail.

Wing loading, as explained by Roskam,5 affects many airplane handling qualities. Table 7 presents a summary ofthose.

High W/S Low W/S

Stall Speed High LowFieldlength (La and To) Long ShortMax. L/D Ratio: High LowRide quality in turbulence: Good BadWeight: Low High

Table 7: Wing Loading effects summarized

5J. Roskam. Airplane Design: Part III: Layout Design of Cockpit, Fuselage, Wing and Empennage : Cutaways and Inboard Profiles. DARcor-poration, 2002.


As a general rule, loadings up to 3 kgm−2 are for gentle flying, 3 kgm−2 to 6 kgm−2 for trainers and above 6 kgm−2

for aerobatics. However, Reynolds6 suggests that wing cubic loading factor, expressed as WCL = WeightArea1.5 , would be

more appropriate to estimate the handling qualities for RC airplanes. Including the linear term (dividing by root ofarea, i.e. some reference length) in the wing loading equation introduces the airplane size as another factor. Thismakes it easier easier to categorize airplanes according to their controllability. Wing cubic loading for the designedaircraft varies from 5.2 kgm−3 to 13.3 kgm−3 for different missions. According to Myers7 the aircraft would categoriseas a sailplane/park flyer for missions 1 and 3, and as an expert sport airplane for mission 2. It was decided that as longas the pilot has enough experience flying the aircraft, handling qualities similar to a sports aircraft are manageable,hence the minimum mass design point was accepted.

Level and Description Average WCL Factor,kgm−3

1: Includes mostly indoor type models and those that can be flown outside in verylight winds, only level with no internal combustion powered planes

2.39

2: Includes mostly backyard type models that can be flown indoors in largervenues and outside in low wind conditions, includes a few internal combustionpowered planes

4.10

3: Includes park flyers, sailplanes, biplanes, 3D planes 5.994: Includes sport types, biplanes, scale, a few 3D planes, pattern, largest level 8.515: Includes advanced sport types, sport scale and sport scale warbirds, sometwins

11.25

6: Includes expert sport types, scale, scale warbirds, twins 14.317: Includes planes for the expert flier only, twins and multi-motor, true scale, war-birds.

17.52

Table 8: RC Airplane general handling quality levels.http://www.theampeer.org/M1-outrunners/M1-outrunners.htm#CWL

After the preliminary CAD model was created, the number of balls to be carried for mission 3 was discussed. Twoconstraints limit this. First, battery capacity to fly the laps required. It was calculated that the battery required formission 2 allows flying another two laps in mission 3. However, the limitations from having the ball cage under thepayload bay are bigger. Taking the wing incidence angle into account, roll angle has to be at least 15◦. Since the ballcage can not be pushed further up front than the nose landing gear, it limits the roll angle unless the landing gearlength is increased. And following the main design goal for the mass to be kept at minimum, it was decided to stickwith 6 balls.

4.4 Estimated aircraft characteristics

Table 9 summarizes the inputs and outputs resulting from the optimization code.

6Francis Reynolds. Wing Cube Loading. Model Builder. Sept. 1989.7Ken Myers. Club Newsletter. Jan. 2014. URL: http://www.theampeer.org/M1-outrunners/M1-outrunners.htm#CWL.


http://www.theampeer.org/M1-outrunners/M1-outrunners.htm#CWL

Inputs:

Used CLmax 1.63Used taper Ratio 0.6Aspect ratio 8Number of balls flown in mission 3 6

Outputs for minimum empty mass configuration:

Masses:Total empty mass, kg 1.449Total mission 2 mass, kg 3.717Wing mass, kg 0.249Boom mass, kg 0.118Motor mass, kg 0.112Battery, kg 0.395Other data:Reynolds number of the wing 223,000Power loading for mission 2, Wkg−1 85Wing area, m2 0.427Wing loading for mission 2, kgm−2 8.7Wing cubic loading for mission 2, kgm−3 13.3Wing span, m 1.849Wing mean aerodynamic chord, m 0.236Root chord, m 0.289Tip chord, m 0.173Power requirements:Takeoff power requirement for mission 2, W 316Best L/D power requirement for mission 2, W 109

Table 9: Optimization code parameters

4.5 Mission performance and Stability analysis

Even though the airplane was designed to fly at maximum speed of 23m s−1 for mission 2, while calculating theof mission 1 the maximum speed was kept the same. The aircraft will experience more trim drag without payload,however this will be compensated by higher dynamic thrust using a higher pitch propeller. An option to raise bothailerons to decrease the lift was discussed, but only after flight tests it could be implied practically. For mission 3 theairplane speed is set to the best L/D speed for maximum cruise performance.

Mission 1 Mission 2 Mission 3

Time, s 230 121 341Energy required, mAh 570 667 465Cl, cruise 0.114 0.291 0.419Stall speed, ms−1 5.77 9.70 6.54Takeoff speed, ms−1 6.347 10.67 7.192Best L/D speed, ms−1 - 13.6 7.81

Table 10: Mission Performance Estimations

It was estimated that 6 laps could be flown for mission 1 in 230 s, which, after comparing numbers from previousreports, was a reasonable number. Mission 2 was estimated to take 121 s, a bit more than a half of the mission 1time. Energy requirement estimation for all the missions was done multiplying the power required for each missionsegment by the time it takes to complete the segment. The results are tabulated in Table 10. The aircraft RAC wascalculated to be 12.78.


4.5.1 Static Stability

Having a big takeoff weight difference in between the missions implies stability issues. After having the preliminaryCAD model of the aircraft ready the CGs for different missions could be estimated. The CG change appeared to bequite radical, if it is to be trimmed with the elevator only. However, since possibility to trim with the battery exists, it ismore useful to keep the CG within a smaller static margin range, which then results in less trim drag and also helpsthe pilot by keeping similar control characteristics of the aircraft.

The most important characteristic in context of static stability is the neutral point location. The aircraft is stable aslong as its CG is in front of the neutral point, which usually depends purely on the aerodynamics of the aircraft. Thedistance between the neutral point and the CG location is called the static margin, which as a first estimation wasapproximated to be 5% MAC. As the static margin grows, the airplane becomes more and more statically stable.Since it is assumed that the elevator servo and control mechanisms are stiff enough, stick-fixed neutral point isconsidered.

Usually, it is best to place the wing so that its AC is quite close to the CG of the aircraft, reducing the tail size required.However, due to structural limitations, it was decided to place the wing in front of the mission 2 payload compartment.This meant that the CG for mission 2 is behind the wing aerodynamic centre, implying that the tailplane is goingto have to provide lift force for stable flight. After assuming the aircraft to be trimmed to the same or more stableconfiguration with the battery in missions 1 and 3, required neutral point position could be estimated and further taildesign could be done.

4.5.2 Tail sizing

0.4 0.6 0.8 1 1.2 1.4 1.6 1.80.1

0.12

0.14

0.16

0.18

0.2

0.22

0.24

Tail Arm Length, m

Tail

Mass, kg

Figure 14: Tail Arm length vs Tail Mass

For this competition, tail design was a very special and unique case. No rudder meant it is required to make surethat the vertical tail is big enough for sufficient directional stability without any active controls. Also, since the CG isquite far behind the wing’s AC, it meant that the horizontal tail is going to have to be big enough to move the neutralpoint of the aircraft further aft. An all moving tail for more control authority and a symmetrical NACA 0010 airfoil foroptimised trim drag for all the missions were selected. Both surfaces were designed with low aspect ratios to reduce


the risk of tail stall.8

To keep the volume coefficients as required, the tail can either be closer to the wing and bigger, or further away, butsmaller. Knowing the densities of the tail boom and the foam that is going to be used for the surfaces, an optimisationwas done for minimum weight. The resulting graph of tail mass with respect to tail arm can be seen in Figure 14. Thetail size was then refined using XFLR5 simulations, which showed a smaller static margin than the optimisation codedid. This most likely was the result of the optimization code not including the pitch up moments coming from the veryhigh camber wing airfoil. Final tail characteristics can be found in Table 11.

Selected inputs:

Horizontal tail volume coefficient. Most influential pa-rameter for the optimisation.

1.09

Vertical tail volume coefficient. Most influential pa-rameter for the optimisation.

0.1

Horizontal tail aspect ratio 5Vertical tail aspect ratio 3.2Horizontal tail taper ratio 0.4Vertical tail taper ratio 0.5CG location for mission 2, m 0.301Optimized tail characteristics:Neutral point location, m 0.315Static margins for mission 2, %MAC 3.8Boom length, m 1.444Horizontal tail area, , m2 0.09Horizontal tail span, m 0.672Horizontal tail root chord, m 0.192Horizontal tail tip chord, m 0.077Vertical tail area, , m2 0.07Vertical tail span, m 0.323Vertical tail root chord, m 0.269Vertical tail tip chord, m 0.135

Table 11: Tail Characteristics

4.5.3 Dynamic stability

The dynamic stability modes were estimated using XFLR5 software while simulating wing and tail only and can befound in Table 12. Without including the fuselage, the simulation resulted in level 1 performance for all modes exceptthe spiral one. Level 2 spiral mode, even dropping lower than level 3 for Mission 2, flying at 10ms−1, was the biggestpredicted handling problem. It was discussed to be a result of a high vertical CG location in relation to the wing. Threesolutions were suggested. First was to flight test the airplane with different vertical tail size and dihedral combinationsto find the most flyable one. Second one was to redesign the wing box so that the wing AC is above the mission 2CG. Third one was to introduce a stabilization system sacrificing the mass as a tradeoff.

8S. Gudmundsson. General Aviation Aircraft Design: Applied Methods and Procedures. Elsevier Science, 2013.


Figure 15: Root Locus plot for the aircraft

Mission 2, 10ms−1 Mission 2, 20ms−1

Mass, kg 3.658 3.658Static margin, %MAC 5.95 5.95CG locations, x, m 0.301 0.301CG locations, z, m 0.036 0.036Phugoid damping ratio 0.326 0.049Phugoid frequency, rad s−1 0.516 0.325Phugoid period, s m 12.182 19.358Short Period damping ratio 1 1Short Period frequency, rad s−1 3.997 10.97Spiral mode time to double, s 1.076 6.016Roll mode time to double, s 0.126 0.036Dutch Roll damping 0.69 0.813Dutch Roll frequency, rad s−1 21.233 31.05Dutch Roll f · d product 14.66 25.25

Table 12: Stability Analysis Results

4.6 Propulsion

In order to ensure a highly optimized propulsion system a multi-step approach is taken as described below.

4.6.1 Methodology

Propulsion selection factors are outlined as follows:


1. Motor and propeller pair is able to provide sufficient thrust for preliminary design specified lift-off and cruiseflight requirements.

2. Propulsion system is able to provide optimal performance for the mission bearing the most points. Consequentlythrough propeller changes and ESC programming the motor selected will be adapted for best performance insecondary missions.

3. Propeller has to be within ground clearance limitations.4. Propulsion system can achieve a speed of 23m s−1 taking into account the overall frame drag generation. This

requirement is imposed in order for the aircraft to be able to withstand the high wind profile in Tucson, AZ. Thisimposes further limitation on propeller size, pitch and motor RPM combinations.

5. The weight of the propulsion system and battery pack combined is to be brought to a minimum, ensuring optimalefficiency of the propulsive system.

Drawing on the above selection the following motor and propeller profiles were established:

Propulsion system profile: A low Kv range, 350Kv to 900Kv, brushless outrunner type motors were selected dueto their:

1. Superior efficiency(a) Lower Kv motors allow higher voltages for a given RPM and power output level. Consequently allowing

lower amps for a given input power, and finally lower amperage permit for smaller diameter wiring andphysically smaller components at a given efficiency, minimizing waste heat. Furthermore, this allowsreduced material costs and weight. Therefore voltage intensive, low kV motors have been found to bemore efficient than high kV ones. However, additional tests are required to ensure high grade motorquality and consequently match it to the specific mission requirements for optimal results. It should alsobe mentioned that the higher voltage requirement impacts negatively on battery cell structure, weightwise, hence a larger range of kV options is used to allow finding the best tradeoff between weight andperformance.

2. Allows direct drive - does not require a gearbox and therefore avoids its negative impact on performance.3. High torque - allows for larger and more efficient propellers, however further testing is to be performed for an

accurate estimate of torque roll performance impact produced by the lower range kV motors selected

Motor Power Range of 315W-600W: Optimization code predicted a minimum required power of 315W for thedesired performance, this was consequently compared to theoretical results from optimization software and statisticaldata from similar motors in order to confirm its consistency. A range of motor maximum output power is presented,adding another dimension to optimization testing , namely optimal efficiency at minimum required thrust complyingwith the requirements set. The spectrum limits are dictated by brushless motors optimal efficiency loading, usuallybetween 40% and 70% of the maximum load, and the minimal difference in weights between the selected models.Range of Carbon fibre propellers with main dimensions range 10 ′′, to 15 ′′, for propeller diameter and 3.7 ′′, to 10 ′′,for propeller pitch:

1. Chosen to accommodate both speed focused and endurance focused missions. Appropriate propellers areto be paired with motors that have suitable RPM and torque parameters in order to satisfy speed and thrustrequirements.

2. Lighter Material - Carbon fiber propellers are significantly lighter than all other commercially available alterna-tives, providing both weight minimization and lowering torque requirements compared to other propellers.

3. Endurance - Carbon fibre provides excellent elasticity and is therefore significantly more resistant to impacts,allowing to improve time efficiency in flight testing phases.

4. Well balanced - Good quality carbon propellers require very little to none weight balancing, which eliminatesthe risk of impacting propeller aerodynamic characteristics during optimization.


Using the profile generated, various pieces of simulation software including OpenProp, eCalc and MotoCalc, andstatistical research on the top contenders of the last four years of DBF, an extensive market research has been donein order to select a range of best suited motors , which is consequently narrowed down to three choices presented inTable 13.

Motor kV Max Contin-uous Power,W

Motor weight, g Max Con-tinuousCurrent, A

ESC

T-MOTOR MN4010 370 450 112 20 Afro HV 20A ESCTurnigy Multistar 3508 640 550 98 30 Hobbywing Platinum

40A-OPTO Pro ESCTurnigy Multistar 4010 485 500 128 26 Hobbywing Platinum

30A-OPTO Pro ESC

Table 13: Table of Motors and ESC considered

Electronic Speed Controller An ESC appropriate for each motor has been chosen with three main parameters inmind:

1. Is able to support the maximum amperage and voltage that the motors are expected to reach without overheat-ing, therefore an extra ten to twenty percent of current capacity have been added to each one. It should benoted that the afro ESC is selected for voltage intensive tests and is therefore not expected to go over 15A

current in any circumstances.2. Is programmable in order to accommodate optimization for both battery type and response times.3. Is as light as possible while complying with first two points.

Fuse selections: A range of fuses will be tested for each motor in order to confirm the characteristics given andensure safety of the propulsion system. The ESC candidates are shown in Table 13.

4.7 Electronics

4.7.1 Battery Selection Factors:

1. The battery must supply adequate power to the propulsion system over the full mission length2. Meet the takeoff maximum power requirement of 315W3. Must have the correct voltage to rotate the propeller at a sufficient angular velocity4. Mass must be under the 2.0 lb (0.9 kg). maximum5. As lightweight as possible6. Battery cell chemistry must be either Nickel Metal hydride (NiMH) or Nickel Cadmium (NiCd)

NiMH cell chemistry offers several significant advantages over NiCad mainly a higher maximum continuous dischargerate (10C) and a larger energy density giving a lower mass for the same energy content.

All performance NiMH cells claim a nominal voltage of 1.2V and a continuous C rating of 10. As, t = 1/C, this meansthe minimum flight time will always be 6 minutes, even if the aircraft is on full power for the entire flight. Most NiMHare also capable of supplying a higher ’burst’ current, usually up to 15C. Selected cells for testing are shown in Table14.

Using a battery with a higher specific energy means the mass of the battery can be reduced while delivering thesame performance.


Battery Type/Manufacturer C rating Capacity,mAh

ContinuousCurrent, A

Burst Cur-rent, A

SpecificEnergy,Whkg−1

Overlander AA high cap pip 10 2000 20 30 88.9Turnigy 2/3A High Power 10 1500 15 22.5 78.3KAN 700 AAA 10 700 7 10.5 60.4Overlander SubC Sport 10 2000 20 30 61.5

Table 14: Electronics cell selection

For mission 2 the takeoff power is 315W, the battery specifications can be calculated.

P = IV

where I is current, V potential difference and P is power.If the maximum current draw is 30A, then a voltage of 10.5Vis required, however most battery packs will drop or ’sag’ under load, sometimes as much as 70%. To account forthis the nominal pack voltage should be around 10.5V/0.7 = 15V.

This is equivalent to a 12 to 14 cell pack, where each cell has a nominal voltage of 1.2V. The lower, 1500mAh batterywould need 20V, or 16-18 cells.

Mission 1 and 3 follow a similar process and will likely use the same battery and propulsion settings.

4.8 Radio System

Transmitter and Receiver requirements:

• At least 4 channel control• Failsafe on all channels• Full range• 2.4GHz or 72MHz band• Telemetry

• Basic timers• Model memory saves• Dual rates and exponential• Trim conditions• Nickel based batteries used to power both

The FrSky Taranis was selected to meet these requirements as well as being considerably cheaper than competitors.The X6R will be used, a 6 channel receiver, to allow telemetry usage during testing and to minimise on board weightcompared to a 8 or 9 channel receiver.

The channels on the receiver are arranged as follows;

1. Ailerons2. Elevator3. Throttle4. Ball release activation

Receiver Battery The receiver battery must supply sufficient current to give the servos full torque as well as supplythe receiver with a steady voltage. A 4 cell (4.8V) 1000mAh Sanyo eneloop NiMH pack was chosen due to its lowmass and ability to supply the required power to the receiver and servos for even more than one flight. Testing was tobe done on the increase in servo torque and speed from using a 6V receiver battery, but the 6V pack may damageother components and was not considered any further.


Servos Because the number of servos is a critical component in the aircraft cost factor NSERVO, it is obvious thatmultiple tasks be carried out by a single servo and to discard any unnecessary servos.

This aircraft requires three servos. A servo to control the elevator providing pitch control, and a larger servo to controlboth the ailerons and steer the nose wheel. The third will command the ball dropping mechanism for mission 3.

The servos supply the torque and movement to drive the control surfaces. Estimates on the required size, andhence torque and speed, of the servo was taken from experience and the relatively slow speed the aircraft will reachcompared to similar sized models.

Obviously a considerably large servo is required if it is to perform multiple functions such as both ailerons and thenose wheel steering. But smaller servos are used for elevator control and release mechanism. A selection of threeservos was purchased to fit the required role and are shown in Table 15.

Servo Mass, g Torque, kg cm−1 Speed (0-60◦), , s

BMS-706 26 4.6 0.13TGY-211DMH 16 1.9 0.12HKSCM12-5 12 1.5 0.18

Table 15: Servo selection

Problems with single servo ailerons Normally, a larger and complex model aircraft would use two servos to driveeach aileron individually. This allows for effects such as differential which can help correct adverse yaw during rollsand add the ability to use the ailerons as flaps, a mix known as flaperons. While flaperons are impossible with a singleservo aileron system, differential can be added to the ailerons using off 90◦control horns on the surfaces, meaningthe ailerons will deflect more up than down.

Servo Linkages Pull-pull wires on nose gear are always in tension, meaning there is never any slack in the system.This can help prevent the aircraft from pulling to one side and help the machine track straight.

Ailerons torque rods are used inside the wing to enable the single servo mounted in the center of the aircraft to controlthe ailerons near the wing tips.

The elevator to servo connection is done using a single push/pull rod, mounted on the upper surface of the elevator.Mounting the servo horn on the top side means the rod is in tension when the elevator is in the ’up’ position, reducingslack in the system and allowing for greater control.

Stabilisation The rules allow for some stabilisation so long as there is no automatic element like auto-level or anyautomatic navigation.

This rules out auto level systems (eg accelerometers), and allows for gyros only which hold orientation. It is thoughta gyro placed on the roll axis could prove valuable to correcting the poor spiral mode characteristics. Further testingis definitely required to calibrate the gyro gain and tune the response to prevent flutter yet still augment the flightstability.


5 Detail Design

5.1 Aircraft Dimensions and Components

Wing

Aerofoil Wortmann FX63-137Span 1.85MAC 0.24Wing Area (S) 0.43Aspect Ratio 8Incidence Angle -3.7Static Margin for mission 2 3.8

Fuselage

Length 0.996Width 0.16Height 0.18

Tail Surfaces

Horizontal VerticalAerofoil NACA 0010Span 0.672 0.323MAC 0.143 0.209Wing Area (S) 0.09 0.07Aspect Ratio 5 3.2Incidence Angle 0 0Tail Arm 1.252 1.242

Table 16: Final Design Dimensions and Aerodynamic Parameters

Controls

ESC Afro HV 20A ESCReceiver FrSky F6RServos BMS-706 (Aileron and Gear Steering)

TGY-211DMH (Elevator)HKSCM12-5 (Sensor Drop)”

Propulsion

Motor T-MOTOR MN4010Mission 1 14x10Mission 2 15x8Mission 3 15x8

Table 17: Aircraft Electronics and Propulsion components


Aircraft Component Weight CG Location

X Y Z(g) (mm) (mm) (mm)

Wings 245 47 -12 -224Fuselage 342 0 19 -298Propeller 55 0 20 37Tail 137 -1 28 -1192Landing Gear 99 0 -147 -327Ball Drop Mechanism 29 2 -11 -523Motor 120 0 20 10Front Electronics (Receiver, aileronservo and low voltage alarm)

52 14 15 -157

Ball Cage 26 0 -99 -409Mission 1: Empty configuration 1585 8 7 -288Battery, position 1 470 0 41 -154Static margin = 11.5%Mission 2: Sensor PackageTransport

3853 3 41 -304

Sensor Package 2268 0 64 -333Battery, position 2 470 0 41 -71Static margin = 4.68%Mission 3: Sensor Drop 1993 6 -13 -2976 x Payload Balls 408 0 90 -336Battery, position 1 470 0 41 -154Static margin = 7.65%

Table 18: Balance Table

5.2 Aerodynamic analysis

After having the final CAD model freezed, further aerodynamic analysis could be done. STAR-CCM+ and XFLR5were used for that.

For all the CFD tests, a ”bullet-shaped” domain extending 8 wing spans (or body lengths) from the actual geometrywas created to simulate half-body with a symmetry plane. Trimmed prism layer mesher was used, resulting in 3-4million cell volume mesh. The mesh had to be refined around all the edges, wing-tip vortices, airfoil leading edgesand smaller details like the ball cage strips.

First, the wing lift predictions were compared for all the available sources. Wind tunnel test results of a 2D airfoil,9

2D XFLR5 test results, 3D finite wing XFLR5 and 3D STAR-CCM+ simulation results were compared and plotted inFigure 16. 3D simulation had a smaller lift curve slope than the 2D, as expected. Also the lift coefficient predicted withSTAR-CCM+ reached a value of 1.66, just a bit higher than what was used in the preliminary design. This confirmedestimations from the preliminary phase to be very close to the actual values.

9Michael S. Selig and Bryan D. McGranahan. In:


Figure 16: Comparison of finite wing performance using different engineering tools

Fuselage, tail, main landing gear and ball cage combination was simulated separately at 0◦ angle of attack overvelocities range of 5m s−1 - 35m s−1. The CFD model for the configuration at reference speed of 14m s−1 calculatedthe parasitic drag coefficient to be 0.0239. Interestingly, the value predicted in the preliminary design stage wasjust 2% different, 0.2364 without the wing. This proved the validity of using empirical drag estimation methods forpreliminary design stage. The refined mesh and the streamlines can be seen in Figure 17.

Figure 17: Refined mesh and streamlines visualized

The drag coefficients of configuration without the wing and the wing separately were added together to get a complete


drag polar. Interference drag between the wing and the fuselage was omitted, since, according to Hoerner,10 dragincrease due to wing placement at the bottom of the fuselage and fuselage incidence angle cancel each other out.Dynamic thrust estimation using an online tool11 for 14x10 propeller was done and is plotted in Figure 18 togetherwith the drag polars.

As can be seen, the actual drag at low speeds for mission 2 is higher than for mission 1. This arises because of the liftinduced drag at higher lift coefficients. For higher speeds the trim drag for mission 1 increases and becomes higherthan mission 2 drag. The preliminary drag polar was higher than the actual one, probably being a result of wronglyestimated planform efficiency. The maximum speeds for both, CFD and preliminary drag models were smaller thanexpected for the selected propeller. Achievable speed of 17.2m s−1 for mission 1 and 14.4m s−1 for mission 2 donot cover the design target of 23m s−1. To cope with the winds, a more powerful motor and battery might have to beconsidered after proving the drag models with flight tests. It also has to be noted that the online tool for dynamic thrustestimation is said to underestimate the propeller performance, hence the actual maximum speeds will be higher.

Figure 18: Drag vs Thrust

Wake structures behind the model were analysed as shown in Figure 19. Filtering the mesh by vorticity, high tur-bulence zones could be identified. The strongest zones appeared behind the first ball cage rib, fuselage end andlanding gear wheel. Weaker zones could be found behind second and last cage ribs, landing gear mount and inter-section between the fuselage and the landing gear. The most problematic spot, the end of the fuselage is going tobe blended with the main boom for the final model. The cage ribs are going to be sanded and streamlined as long asthe structural strength is conserved and landing gear fairing might be considered if drag reduction will compensatefor the mass addition. One more ”red” zone can be seen at the joint of the main fuselage compartment and the frontfairing. Separation occurs there and sanding the frontal fairing to a more rounded shape is suggested to decrease

10Hoerner, Fluid-dynamic drag: practical information on aerodynamic drag and hydrodynamic resistance.11Gabriel Staples. Article. Apr. 2014. URL: http://electricrcaircraftguy.blogspot.co.uk/2013/09/propeller-static-dynamic-

thrust-equation.html.


http://electricrcaircraftguy.blogspot.co.uk/2013/09/propeller-static-dynamic-thrust-equation.html

http://electricrcaircraftguy.blogspot.co.uk/2013/09/propeller-static-dynamic-thrust-equation.html

the adverse pressure gradient. It is to be noted that the propeller wake effects were not included in the simulationbecause of the complexity of the model required.

Figure 19: Wake structure behind the fuselage, landing gear, ball cage and the tail

5.3 Structural characteristics

The structure sub team’s design goals were to minimize the structural weight of the aircraft, ensuring that the structurewill provide adequate strength and rigidity for high wing loading during turns.

5.3.1 Load Paths

The main structural components of the aircraft are several carbon fiber tubes. A single 12mm outer diameter tuberuns along the span of each of the wing pieces at 25% chord. To ensure torsional rigidity of the wing, a 4mm outerdiameter tube runs along the span of each of the wing pieces at 75% chord. A 12mm outer diameter carbon fibertube runs through the foam nose cone, the carbon fiber bulkheads and the foam aft part of fuselage. Those carbonfiber tubes and the connections between them form the main load paths. The load paths are presented in Figure 20.


Figure 20: Load Paths

5.3.2 Structural Analysis

The main structural components of the aircraft were analyzed. An assumption was made that the carbon fiber tubesalong the wing pieces and the fuselage carry all of the structural loading. The empty aircraft undergoes a load factorof five during turns. The loading applied to the 25% chord spars is shown in Figure 21. An assumption was madethat the 25% chord spar of the wing features distributed loading due to the lift and point loading due to the weight ofthe aircraft.

Figure 21: 25% chord spar loading

Based on the loads, shear and moment diagrams were created. This analysis does not account for spar deflection,which may cause other structural components to fail well before the spar fails. A diagram is shown in Figure 22.


Figure 22: Loading, shear and moment for wing 25% chord spar

The main load carrying structure of the fuselage is a carbon boom. As with the rest of the components, the intentionwas to identify the lightest, yet sufficiently stiff and strong boom. The two available suitable options were 12mm

pultruded and rolled carbon fibre tubes. Maximum load during 5.6G turn has been modelled using ABAQUS FEA.Since the supplier provided a range of values for tensile moduli, it was decided to use a safety factor of 1.5 rather than1.2, that was used during the actual testing, in order to make sure that the selected boom satisfies the performancerequirements. Hence 8.4G loading was used for modelling. The pultruded boom was predicted to deflect significantly,and tensile modulus exceeds the maximum strength of the material, which immediately ruled it out as an option.Rolled boom, however, while being about 10% heavier, would experience a much smaller deflection ,as can be seenin Figure 23. The stress is also within maximum tensile strength. It was thus decided to acquire a rolled tube andconduct live tests in similar conditions.


Figure 23: Main boom deflection under 8.4G MTOW load factor

5.4 Aircraft Systems Design, Component Selection and Integration

5.4.1 Fuselage

The fuselage was designed around the internal payload bay. It had to have a small frontal area and to be as light aspossible. It consists of a carbon fiber tube of 12mm outer diameter running along the entire length of the fuselage,carbon fiber motor mount, carbon fiber bulkheads and flat carbon fiber spars connecting the bulkheads. A carbon fiberplate secured between the motor mount and the first bulkhead is used as a mounting plate for electrical components.The plate is also secured to the fuselage spar. As an effort to minimise weight the nose cone and the empennagewere not made by placing additional bulkheads and spars but were made entirely from foam. The foam empennageand nose cone are lighter but still contribute to the structural rigidity of the fuselage. The nose cone is secured to themotor mount, the first fuselage bulkhead, the mounting plate for electrical components and the fuselage spar. Theempennage is secured to the third fuselage bulkhead and the fuselage spar. A 12mm ID spar which is perpendicularto the fuselage spar is located between the first and the second fuselage bulkheads. The 25% chord spars fromeach wing piece slide into this spar. The remaining space between the bulkheads above the spar is filled with foam.Another carbon fiber plate between the second and third bulkhead is used for mounting payloads. The third bulkheadfeatures the mount for the main landing gear. Renderings of the fuselage are shown in Figure 24.


Figure 24: Fuselage

5.4.2 Internal payload bay

The internal payload bay is situated between the fuselage second and third bulkhead. A rectangular carbon fiber plateis placed between the bulkheads and onto the fuselage spar. The plate serves as a mounting point for the payload.The payload bay is accessible by a foam cover. Two carbon fiber plates are glued to the foam cover that are insertedinto openings in the second bulkhead. A simple locking mechanism at the back of the foam cover prevents movementof the cover during flight. It is doing this by interacting with the third bulkhead. This mechanism is comprised of a bentrod which latches onto an opening in the third bulkhead. Internal payload bay is depicted in Figure 25. The payloadand battery are secured with velcro straps.

Figure 25: Internal payload bay

5.4.3 External ball cage

The external ball cage is situated below the fuselage and consists of three flat carbon fiber brackets and two flatcarbon fiber spars as seen in Figure 26. The brackets are designed in such a way that the external payload bay canbe detached from the fuselage. The payload bay is specifically designed to carry a payload in the form of balls whichcan be dropped with the help of the ball drop system.


Figure 26: External Payload Bay

5.4.4 Wing

The wing is made up of two pieces. Each piece consists of carbon fiber ribs, two carbon fiber spars and a foamleading edge, trailing edge and aileron. A 12mm outer diameter spar is present at 25% chord and a 4mm outerdiameter spar at 75% chord. A 10mm external diameter spar of 300mm length is inserted 150mm into one of the 25%chord spars of the wing pieces. The 25% chord spar of each wing piece is inserted in the wing attachment tube. Thewing can be seen in Figure 27.

Figure 27: Wing

5.4.5 Wing attachment

The wing attachment consists of a carbon fiber tube featuring 12mm internal diameter placed below the fuselage sparand between the first and second fuselage bulkheads as depicted in Figure 28. The wing attachment spar featuresa small hole. As mentioned before a carbon fiber tube of 10mm external diameter and length of 300mmis inserted150mm into the 25% chord spar of one of the wing pieces. This spar is secured to the 25% chord spar and featuresa small hole. The 25% chord of the other wing piece also features a small hole. When the two 25% chord spars ofeach wing piece are inserted into the wing attachment a three spar connection is formed. With the holes aligned apin is placed, which secures the two wing pieces. With this pin the wing can easily be attached or detached from thefuselage.


Figure 28: Wing attachment

5.4.6 Tail

The horizontal and vertical tail are made out of foam, and they feature a carbon fiber spar of 10mm outer diameterwhich makes them more stiff and resistant to damage. Illustration is provided in Figure 29.

Figure 29: Tail

5.4.7 Tail attachment

The tail attachment consists of two carbon fiber tubes with 12mm outer diameter and carbon fiber rod with 10mm

outer diameter. The horizontal carbon fiber tube of 10mm outer diameter is perpendicular to the carbon fiber rod. Thevertical carbon fiber tube of 12mm outer diameter is perpendicular to both the horizontal carbon fiber tube and thecarbon fiber rod. The tail spar of 10mm outer diameter passes through the horizontal carbon fiber tube of 12mm outerdiameter. This allows the horizontal tail to pivot about its spar acting as a stabilator. The vertical tail spar of 10mm

outer diameter is attached to the vertical carbon fiber tube of 12mm outer diameter. Tail attachment is rendered inFigure 30.


Figure 30: Tail Attachment

5.4.8 Fuselage aft section

This component provides the aft aerodynamic profile of the fuselage. It is designed to be lightweight. Instead ofmaking the shape of the empennage with carbon fiber bulkheads, foam was used. The foam empennage featureslightening holes and the fuselage spar runs through it. The empennage is secured to the fuselage spar and the thirdfuselage bulkhead. The servo for the horizontal tail is located at the very end of the foam empennage. The servois attached to mounts made out out of plastic which are attached to the fuselage boom. This part also houses thesensor drop mechanism. This part of the fuselage is shown in Figure 31.

Figure 31: Aft part of the fuselage

5.4.9 Nose cone

The nose cone as presented in Figure 32 provides the forward aerodynamic profile of the fuselage. It is designed tobe lightweight, so it features lightening holes just like the empennage. The nose cone is secured to the fuselage spar,the motor mount, the first fuselage bulkhead and the carbon fiber plate used for mounting electrical components. Thetop aft section of the nose cone is removable offering quick access to the electrical components located inside.


Figure 32: Nose Cone

5.5 Payloads Systems Design, Component Selection and Integration

5.5.1 Ball drop system

The ball drop system is made of one carbon fiber tube of 14mm external diameter and two carbon fiber plates(paddles) of 4mm thickness attached to it. There is an angle between the paddles. The carbon fiber tube featuresa control horn which is connected via a push-pull rod to the control horn of the servo driving the ball drop system.As the control horn of the servo moves left to right the carbon fiber tube rotates and so do the the carbon paddlesattached to the tube. As the paddles rotate the aft paddle gets out of the way of the last ball in the external payloadbay and the forward paddle gets in the way of the next ball. This prevents additional balls from falling. After the lastball falls the paddles are rotated which allows the next ball to move in place. Airflow pushes the balls backwards. Balldrop system is depicted in Figure 33.

Figure 33: Ball Drop system


184872.76

166.676.56

56622.28380

14.96 241.68

9.52

89.883.54

199.097.84

70927.91

37414.72

172.926.81

287.2611.31

1606.30

1164.57

1204.72

33913.35

THE REST OF THE DISTANCES FROM

PROFILE TO PROFILE

UNTIL THE END OF THE

WING Is 120mm

32012.60

27010.63

153260.31

57522.64

99939.33

1807.09

42916.89

SCALE 1:14

TITLE:

SIZE

BUNLESS OTHERWISE SPECIFIED DIMENSIONS ARE GIVEN IN MM AND THEN IN INCHES

SCALE: 1:10(Unless specified)

Drawing number№1

SHEET 1 OF 4

Cessna/Raytheon Missile Systems Student Design/Build/Fly competition

AIRCRAFT 3-VIEW DRAWING

SolidWorks Student Edition. For Academic Use Only.

2

139

10

4

8

56

7

14

12

1315

16

18

19

2021

22

23

24

11

17

Part Material Quantity1 Main Spar Carbon Fibre 12 Fuselage Foam and Carbon

Fibre 1

3 Fuselage Forward Firing Crafting Foam 1

4 Wing Foam Crafting Foam 25 Main Wing Spar Carbon Fibre 26 WingSpar at 75%

Chord Carbon Fibre 27 Aileron Torque Rod 28 Wing Rib 169 Propeller 1

10 Propeller Spinner

Steel Carbon Fibre Carbon Fibre

2014 Aluminium Alloy 1

11 Electrical Mounting Plate Carbon Fibre 1

12 Nose Wheel 113 Main Wheel 214 Front Landing Gear 115 Main Landing Gear

Plastic and Rubber Plastic and Rubber

Steel

Carbon Fibre 1

16 Payload Holder andRelease Mechanism Carbon Fibre 1

17Fuselage Locking

Mechanism for Main Payload

Steel 1

18 Fuselage Aft Firing Crafting Foam 119 Tail Horizontal Spar Carbon Fibre 120 Main Spar to Tail

Connector Plastic 121 Vertical Tail Spar Carbon Fibre 122 Vertical Stabiliser Crafting Foam 1

23 Horizontal Stabiliser Profile Carbon Fibre 2

24 Horizontal Stabiliser Crafting Foam 2

TITLE:

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Drawing number№2

SHEET 2 OF 4


AIRCRAFT Structural Arrangement


G

Payload 2

Payload 1

25710.12

E

Payload 2Payload 2 housing

DETAIL ESCALE 1 : 2

Release paddle 1 Release paddle 2

DETAIL GSCALE 1 : 1

Fuselage LockingMechanism for MainPayload

K

DETAIL K(Payload 1 Housing)SCALE 1 : 6

Location of motor battery for mission 1 and 3

NOTE: Top Foam Hatch has been ommited for clarity

TITLE:

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Drawing number№3

SHEET 3 OF 4


Payload Layout


E

G

SCALE 1:15TOP VIEW

DETAIL ESCALE 1 : 2

PROPELLER

MOTOR

PROP SPINNER

MOTOR BATTERY

RECEIVER

FRONT LANDING GEAR & AILERONS SERVO

LOW VOLTAGE ALARM

SERVOS BATTERY PACK

DETAIL GSCALE 1 : 1

ELEVATOR SERVO

SERVO MOUNT

H

ELEVATOR SERVORELEASE MECHANISM SERVO

FRONT LANDING GEAR MAIN LANDING GEAR

DETAIL HSCALE 1 : 1

RELEASE MECHANISM SERVO

TITLE:

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SCALE: 1:10(unless specified)

Drawing number№4

SHEET 4 OF 4


SYSTEMS LAYOUT


6 Manufacturing Plan and Processes

6.1 Manufacturing Process Selection

The manufacturing process selected for fabrication plays an important role in aircraft weight and overall flight score.There are a lot of manufacturing methods for radio controlled aircraft. The most common are outlined here.

Balsa build-up -the most widely used method, parts from balsa are easy to manufacture by hand, semi-complexshapes can be made with thin sheets of balsa.

Foam - often heavier than a balsa build up, foam parts can be manufactured very easily with a hotwire machine.Foam parts are excellent for making complex shapes such as fuselage lofted sections or wing leading or trailingedges.

Carbon fiber - very rigid, carbon fiber parts can be manufactured relatively easily out of pre cured carbon fiber sheetsusing a CNC machine. Other carbon fiber products such as tubes and rods are readily available.

6.2 Subsystems manufacturing

In the following sections, the manufacturing processes used for the wing, fuselage, empennage, nose cone and othercomponents are documented.

6.2.1 Wing manufacturing

The wing is composed of CNC cut carbon ribs, two carbon fiber spars, and foam. The ribs are aligned along thespars at their respective locations. Epoxy is used to glue the carbon fiber ribs to the carbon fiber spars. Foam blocksare inserted at the leading edge between each pair of ribs. A manual hotwire cutting tool is then used to shape thefoam blocks. The carbon fiber ribs are used as a hotwire cutting templates. The same is done for the trailing edge.The foam provides surface to which the covering film is attached, as well as extra rigidity at low mass penalty.

6.2.2 Fuselage manufacturing

The fuselage is composed of CNC cut carbon fiber parts, carbon fiber boom and foam. CNC-cut carbon fiber bulk-heads are aligned along the main boom at their respective locations. Epoxy is used to glue the carbon fiber parts.The wing attachment carbon fiber spar is placed between the first two bulkheads below the fuselage boom. Payloadand electrical carbon fiber mounting plates are placed directly on the main boom. The aft fuselage section and nosecone are aligned with the bulkheads and glued to them and the fuselage boom. Foam provides surface to which thefilm is attached, as well as stiffens the structure.

6.2.3 Aft fuselage section manufacturing

Profiles which form the shape of the empennage are cut from thick cardboard which are then used as hotwire cuttingtemplates. A manual hotwire cutting tool is then used to cut the lofted shape between the two profiles. The lofted cutfoam parts are then aligned along the fuselage boom and glued with epoxy to it and to the bulkheads.

6.2.4 Nose cone manufacturing

Same technique when manufacturing the aft section is used to manufacture the nose cone. A section of the nosecone is cut with a manual hotwire tool which acts as an access hatch to the electrical equipment. After all fuselagesections are combined, they are covered with film.


6.3 Landing gear manufacturing

The landing gear shape is cut from thick cardboard which is then used as a profile for hotwire cutting. A manualhotwire cutting tool is used to cut the landing gear mold from foam. Carbon cloth is layered on mold multiple times asshown in Figure 34 and covered with epoxy, which is then left to dry for at least 24 hours. It is then sanded down toremove imperfections and acquire the desired shape and size.

Figure 34: Landing Gear manufacturing

6.4 Schedule

19 26 02 09 16 23 02 09 16 23 30 06

Prototype 1

Carbon fibre & Foam cutting Wing assembly

Main Section assembly

Landing Gear assembly

Insert Motor, Servos & Wiring

Prototype 2

Carbon fibre & Foam cutting

Wing assembly




Final Design

Carbon fibre & Foam cutting

Wing assembly




April

Activity

January February March Legend

/

Planned

Actual

Due date

Current

Table 19: Manufacturing Schedule for the project

Every effort has been made to follow the manufacturing plan, however for the reasons described above there wasonly a limited amount the team could achieve. This, however, was sufficient to prove the ability to manufacture allcomponents required and confirm that further plans are possible to meet.


7 Testing Plan

Components testing is an essential part of aircraft design. It allows the verification of predicted performance andprovides data on actual performance. Based on tests performance can either be increased or decreased to optimizethe design according to project specification.

7.1 Testing Objectives

The objectives of aircraft testing were to verify that the design can meet all the rules and requirements of the compe-tition and at the same time accumulate the highest possible competition score. The general goals were as follows

• Determine the most suitable propeller and motor combination– Carried out on a test rig to identify propeller-motor combinations that provide greatest static and dynamic

thrusts• Evaluate individual cell and complete battery performance

– Maximal discharge rate and actual performance under load such as voltage drop and capacity reductionfrom the rated value. Verify sufficient cooling in flight by stress testing at worst expected

• Ensure sufficient structural integrity in all expected flight conditions– Wing and main boom bending test at simulated 6G load and landing gear at 3G

• Test handling characteristics and landing and takeoff performance– Test flight schedule at different payloads and flight conditions will be carried out to verify and adjust sizing

and configuration if required. Ground handling and stability needs to be verified as well.• Ensure the reliability of the sensor drop mechanism in various flight situations

7.2 Propulsion Test

The next step in the selection process is to test the motors static thrust while utilizing different propellers in orderto determine the best combination.This will be done using a carefully calibrated load cells and a power supply unit.The voltage and power output will be kept constant in order to satisfy battery limitations. Optimal ESC and fuseselection for the given motor will be implemented in the corresponding test circuit, along with insulation tape, andbullet connectors for safety.

As stated above main selection will be done based on optimal characteristics for the mission bearing the largestamount of points , therefore a static thrust data table will be created for each appropriate motor and propeller set.

A secondary propeller will also be chosen for the selected motor with respect to other missions optimization.

The selected pairs will consequently be tested in a circuit with batteries instead of PSU, in order to confirm that theresults are viable and apply fine tuning where needed .

Finally ESCs will be reprogrammed in order to achieve best results with the battery circuits.

Practical Testing Sequence: After the lab testing sequence is complete the propulsion system will be used in aseries of flight tests to further confirm the selection and fine tune it where required.

7.3 Electronics Test

Battery testing will be carried out with the use of a custom built, computer controlled, constant current discharger thatuses modular load boards consisting of regulators that provide a set amount of power to banks of resistors. Eachload board carries a pair of thermistors for monitoring board temperatures to provide a layer of safety during extended


and strenuous testing, as well as 2 relays, each of which connects one load circuit to the battery, allowing for finercontrol of the load.

These ’load boards’ are connected to both the battery and control board, which detects the load boards currentlyconnected, controls the relays to control the load as well as monitoring the load board temperatures. In addition, thecontrol board also monitors cell voltage and temperature for logging. The system allows for battery testing to mirrorreal world flight profiles, as it is possible to upload a test profile to the control board processor which can vary theload during testing to simulate take-off power, turning, climb and landing loads on the battery.

7.4 Structural Testing

7.4.1 Wing spar and fuselage boom testing

For the wing spar testing, the objective is to simulate a load during 5G turn with a 1.2 safety factor which results in6G total load. The loads are at MTOW. The lift distribution will be represented as a point load acting at the meanaerodynamic chord. Water bottles will be used as a load. The deflection of the spar at the wing tip at different loadswill be measured. Similar approach will be used with the fuselage boom. The boom will be suspended from twopoints, wing and tail aerodynamic centers, and load applied at aircraft center of gravity. Deflections will be measuredat the maximum deflection point of the boom.

7.4.2 Landing gear testing

For the landing gear testing the objective is to simulate a 3G load representing impact during landing. The landinggear is secured in such way that lateral motion is restricted. The load is applied to the landing gear at the place wherethe landing gear is attached to the fuselage. The deflection of the landing gear is then measured.

7.4.3 Sensor Drop Test

Sensor drop mechanism will be tested as a standalone unit to investigate the operating limits of the system. Theball cage will be tilted at different angles from horizontal to model the force created by the air flow in a variety ofairspeeds. Also the cage is to be tilted from vertical to test sufficient performance even at high bank angles that couldbe encountered.

7.4.4 Flight Test

Flight testing is to be conducted following a carefully planned schedule to slowly expand the flight envelope of theaircraft. Testing was divided into following parts

1. Taxi Testing• The aircraft will be taxied at gradually increasing velocities at different payloads and the aircraft behaviour

will be observed. This is to be continued until 80% of the planned takeoff velocity is achieved. This willensure tolerable ground handling qualities such as turning and stability.

2. Empty Flight Tests• The aircraft will be flown in ferry condition. First flight was planned to only include a small hop to at a low

altitude to verify general flight behaviour and controllability. The flight envelope will be slowly expandeduntil maximum speed and load factor are reached.

3. Flight with variable payloads• After successful empty flight test programme the aircraft will be loaded with variable payload weights and

CG positions to verify flight handling characteristics in different payload configurations before MTOW testcan be carried out.


4. Maximum Payload Test• Aircraft to be loaded to MTOW to measure landing and takeoff distances as well as tolerance to maximum

planned load factor in flight.5. Sensor Drop Testing

• At the last stage of the flight test programme the sensor drop mechanism will be tested and verified towork reliably in all expected drop conditions.

7.4.5 Schedule

03 10 17 24 1 8 15 22 29 5 12 19 26 02 09 16 23 02 09 16 23 30 06

Propulsion test

Component test

Wing Spar test

Fuselage Boom test

Battery cell test

Landing gear test

Drop sensor test

Prototype 1 taxi&flight

Prototype 2 taxi&flight

Final Design taxi&flight

Current Date

Legend

PlannedActual

Due date /

Activity

November December January February March April

Winter

Exam

Figure 35: Testing Schedule

Despite having to shift the actual to the right, all major components were successfully tested before the report sub-mission.


7.5 Flight Checklist

Inspection Item TaskInitial Checks

Aircraft Visually verify that the aircraft in general is in normal condition withoutnoticeable cracks, missing parts or unsecured components

�

Wing No fractures present and wing is firmly in place. Control surfaces moveeasily.

�

Empennage All control surfaces move easily. No cracks or loose components. �

Payload Payload is properly secured and the payload is correctly located. Ifballs are loaded verify smooth action of the drop mechanism.

�

Batteries Make sure battery contacts and leads are in good condition for propul-sion, receiver and transmitter batteries.

�

No visible cracks and wrapping is intact. �

Batteries are fully charged. �

Battery positions are appropriate for the payload carried. �

Landing Gear Firmly in place, no fractures. Tyres spin easily and front gear turns innormal limits.

�

Motor Mount is free of fractures. Securely in place. �

Propeller Careful visual inspection of the propeller. No cracks or fractures.Mounting is appropriate and firm.

�

Center of Gravity Ensure CoG matches with the loaded payload. Max aft position al-lowed is 0.3m from the front tip of the main boom.

�

Pre Flight ChecksFlight Log Fill in Flight Log details �

Batteries Connect batteries and then connect fuse. Make sure the connectionsare good.

�

Radio controls Transmitter on first, Receiver on second. �

Control surfacesand front landinggear

Verify that maximum deflections are nominal and front gear turns withaileron inputs (ailerons full left-right, elevator full up-down), responseto control inputs is consistent. Small slack. Range check.

�

Motor While holding the aircraft from behind, throttle up to full power swiftly.Note response. No excessive vibrations or unusual noise.

�

Post Flight ChecksFuse and Batter-ies

Unplug the fuse and batteries �

Radio controls Receiver off first, transmitter second �

Aircraft Visually verify that the aircraft is in normal condition without noticeablecracks, missing parts or unsecured components.

�

Flight Log Ensure flight is properly documented. Report any discrepancies. �


8 Performance Results

8.1 Structural performance

8.1.1 Wing spar and fuselage boom performance

(a) Wing Spar (b) Main Boom

Figure 36: Structural Testing

The maximum deflection at the tip was measured and the general behaviour of the spar was carefully observed.Significant amount of bending resulted at maximum load factor but the spar remained intact. For the prototype, it wasdesired to have a larger safety factor. For the final production model, destructive testing will be performed to moreaggressively optimise the weight of the aircraft.

Deflection measurement 1, mm 2, mm Average, mm

2G 24.6 24.4 24.54G 35.4 35.3 35.356G 50.1 50.3 50.2

Table 20: Fuselage boom deflections under different loads

Deflection measurement 1, mm 2, mm Average, mm

2G 24.2 24.5 24.353G 50.6 50.2 50.44G 74.5 74.9 74.75G 105.3 104.7 1056G 141.7 140.1 140.9

Table 21: Wing spar deflections at wingtip under different loads

8.1.2 Landing gear performance

The landing gear successfully withstood the 3G load. The deflection at the centre of the gear was just about 1.2cm.There was no permanent twist or bending after unloading, and it was concluded that a lighter landing gear will bemanufactured for the next prototype in order to optimize mass further.

8.2 Electronics Test Results

The battery cells tested so far are the Turnigy 2/3A 1500mAh and the KAN 700mAh. Both of these cells were testedat 5C and 10C, to compare the performance at high current draws. For the purposes of easy comparison C rating


and % Capacity were used to give nondimensionalized results. Results are plotted in Figure 37.

Figure 37: Battery Discharge test results

From initial constant current tests it is apparent that a higher discharge rate causes the battery voltage to sag anddrop below the stated value for all batteries. The KAN700 cells perform well initially by holding a higher voltage thanthe Turnigy cells, but they drop off steadily. At 10C, the KAN cell fails to deliver its full stated capacity and dropsbelow the minimum cell voltage. The Turnigy 2/3A 1500mAh cells perform well, delivering the full stated capacity atboth C levels and holding the voltage in a relatively flat level for most of the tests. There was only a significant dropoff after 80% capacity, but this is not important as the aircraft has short flight times and will likely only use a portionof the battery capacity. Because of the ability to supply a high current for a longer time period, the Turnigy 2/3AA1500mAh cells were selected to be used in the battery.

8.3 Propulsion Test Results

The static test was conducted in the wind tunnel as can be seen in Figure 38. The wind tunnel was not running asthe diameter of the propellers tested exceeded the size of the wind tunnel outlet.


Figure 38: Propulsion wind tunnel test setup

After extensive testing on a large array of motor-propeller pairs, a database of set specifications was created. Theresults in Table 22 were used to estimate mission efficiency as a function of minimum thrust required depending onmission specification, estimated maximum speed , and optimal battery weight. Drawing on these the T-motor MN4010was found to be the most suitable option complemented by a 15x8 carbon propeller for sufficient thrust capabilitiesrequired for missions 2 and 3, and a 14x10 propeller providing additional speed at the expense of static thrust forMission 1. The preselected ESC for this motor (Afro HV 20A) was also confirmed to work well in this configurationwith no signs of overheating. Results for the configuration can be seen in the following table:

Power Actual, W Prop Voltage, V thrust, g RPM current, A

317.2 15x8 25.9 1903 5178 12.25316.4 14x10 25.9 1825 5784 12.22

Table 22: Static Thrust test data

The thrust data obtained agrees well with power requirements produced by the mass optimization code, meaningthat the propulsion matches the design and the aircraft is expected to perform as required.

8.4 Flight Test Results

To team’s great disappointment and despite its best effort, it was impossible to produce a flying prototype on timefor various reasons described above. To compensate for that at least partially, an additional analysis was made toestimate the aircraft performance using CFD for drag estimation and FEA for structures.


Documents

AIAA DBF '14/'15 Report