AIAA DBF '15/'16 Report

School of Engineering Department of Aerospace Sciences

2015/2016 AIAA Design Build Fly February 22 2016

Contents 1 Executive Summary

2

2 Management Summary

3

5.4 Aircraft Mission Performance . . . . . . . . . 32 5.5 Aerodynamic Analysis . . . . . . . . . . . . . 33

2.1 Milestones Chart . . . . . . . . . . . . . . . .

3 Conceptual Design

4 4

5.6 Structural Characteristics . . . . . . . . . . . 34 5.6.1 Load Paths . . . . . . . . . . . . . . . 34 5.6.2 Structural Analysis . . . . . . . . . . . 35 5.7 Aircraft Systems Design, Selection and

3.1 Design Constraints . . . . . . . . . . . . . . .

4

3.2 Scoring Outline . . . . . . . . . . . . . . . . .

5

Integration . . . . . . . . . . . . . . . . . . . . 36

3.2.1 Rated Aircraft Cost . . . . . . . . . . .

5

5.7.1 Production Aircraft . . . . . . . . . . . 36

3.2.2 Total Mission Score . . . . . . . . . .

6

5.7.2 Manufacturing Support Aircraft . . . . 38

3.3 Total Score and Sensitivity . . . . . . . . . . .

7

5.8 Rated Aircraft Cost . . . . . . . . . . . . . . . 40

3.4 Configuration Selection . . . . . . . . . . . .

9

6 Manufacturing Plan

49

12

6.1 Process . . . . . . . . . . . . . . . . . . . . . 49

4.1 Design Methodology . . . . . . . . . . . . . . 12

6.2 Aircraft Subsystems . . . . . . . . . . . . . . 49

4.1.1 Airfoil Selection . . . . . . . . . . . . . 12 4.2 Parameter Selection and Estimation . . . . . 13

6.2.1 Wing . . . . . . . . . . . . . . . . . . . 49

4.2.1 Drag Estimation . . . . . . . . . . . . 13

6.2.3 Tail . . . . . . . . . . . . . . . . . . . . 50

4.2.2 Mass Estimation . . . . . . . . . . . . 15

6.2.4 Landing Gear . . . . . . . . . . . . . . 50

4.3 Constraint Analysis . . . . . . . . . . . . . . . 16

6.3 Schedule . . . . . . . . . . . . . . . . . . . . 50

4 Preliminary Design

6.2.2 Fuselage . . . . . . . . . . . . . . . . 49

4.4 Mission Constraints . . . . . . . . . . . . . . 17 4.4.1 Production Aircraft . . . . . . . . . . . 17

7 Testing Plan

50

4.4.2 Manufacturing Support Aircraft . . . . 18

7.1 Propulsion Testing . . . . . . . . . . . . . . . 51

4.5 Trade-off Studies . . . . . . . . . . . . . . . . 19

7.2 Battery Testing . . . . . . . . . . . . . . . . . 51

4.6 Estimated Parameters . . . . . . . . . . . . . 21

7.3 Structural Testing . . . . . . . . . . . . . . . . 52

4.7 Stability and Trim . . . . . . . . . . . . . . . . 22

7.3.1 Wing and Fuselage Spar Testing . . . 52

4.7.1 Static Stability . . . . . . . . . . . . . 23

7.3.2 Landing Gear Testing . . . . . . . . . 52

4.7.2 Dynamic Stability . . . . . . . . . . . . 25

7.4 Flight Test . . . . . . . . . . . . . . . . . . . . 52

4.8 Propulsion . . . . . . . . . . . . . . . . . . . . 26

7.5 Testing Schedule . . . . . . . . . . . . . . . . 52

4.8.1 Motor and Propeller Selection . . . . . 27

7.5.1 Checklist . . . . . . . . . . . . . . . . 53

4.8.2 ESC Selection . . . . . . . . . . . . . 28 4.8.3 Battery Selection . . . . . . . . . . . . 28 4.9 Electronics . . . . . . . . . . . . . . . . . . . 29 4.9.1 Servo Selection

. . . . . . . . . . . . 29

8 Performance Results 8.1 Propulsion Testing Results

53 . . . . . . . . . . 53

8.2 Battery Testing Results . . . . . . . . . . . . 55 8.3 Structural Testing Results . . . . . . . . . . . 55

30

8.3.1 Spars . . . . . . . . . . . . . . . . . . 55

. . . . . . . . . . . . . . 30

8.3.2 Landing Gear Testing Results . . . . . 56

5.2 Aircraft Weights and Balance . . . . . . . . . 31

8.4 Flight Testing . . . . . . . . . . . . . . . . . . 57

5 Detail Design 5.1 Aircraft Dimensions

5.3 Aircraft Flight Performance Parameters

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. . . 32

8.4.1 Results . . . . . . . . . . . . . . . . . 57

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1

Executive Summary

The objective of this report is to describe the design, manufacturing and testing conducted by the University of Glasgow AIAA Design, Build and Fly team to create aircraft that are highly competitive yet compliant with the competition rules. The aim this year is to create two radio controlled aircraft, the Production Aircraft (PA), that can be split into sub-assemblies, and the Manufacturing Support Aircraft (MSA), that is able to carry the PA sub-assemblies. The aircraft need to perform the following flight missions: 1. Mission 1 The MSA must takeoff, fly three laps following a designated course within five minutes without a payload and land.

2. Mission 2 The MSA must be loaded with one or more PA sub-assemblies, takeoff, then fly one lap following the course, land and taxi to a designated payload loading area. The procedure is to be repeated as many times as needed to transport all the PA sub-assemblies. The mission has to be completed within 10 minutes.

3. Mission 3 The PA must be able to fly three laps within 5 minutes around the course with a 32 oz Gatorade bottle (1020 g actual mass) stowed internally and land.

Upon successful completion of Mission 2, the team may attempt the bonus ground mission where the ground crew must assemble the PA from its sub-assemblies and install its payload within 2 minutes. Maximum takeoff distance is 100 ft (30 m) for all the flight missions.

The total score for the team is obtained from three components: Total Mission Score, Written Report Score and Rated Aircraft Cost (RAC). The total mission score is a multiplication of all flight mission scores plus the bonus ground mission.

The RAC includes the empty weights of the MSA and PA, the MSA and PA battery weights and the number of sub-assemblies PA is split into for Mission 2. Achieving the lowest RAC possible is crucial as it divides the product of the Written Report Score and the Total Mission Score.

From these requirements it was determined through competition score analysis that empty weights and the number of PA sub-assemblies were the two driving factors in the design, followed by the number of required flights in Mission 2 and the battery weights of the MSA and PA. It was evaluated that the team would not benefit carrying the PA as a single sub-assembly in Mission 2 due to the difficult design and manufacturing of the sufficiently efficient MSA in that case. The PA carried in 2 sub-assemblies over 2 flights in Mission 2 was determined to be the highest scoring option that allowed more efficient use of the MSA structure. This is based on the fact that splitting the PA into two sub-assemblies allows the team to effectively halve the payload requirement of MSA for each flight in Mission 2 reducing the MSA empty weight. The reduced empty weight of the MSA can outweigh the effect of more PA sub-assemblies in RAC calculation.

The MSA final design is a high wing monoplane with an empty weight of 0.997 kg and a maximum flight speed of 20 m s−1 .

The main construction materials were carbon fiber, balsa and foam. The aircraft is predicted to be able to complete 3 laps in 119 s for Mission 1, and transport all the PA sub-assemblies in Mission 2 in a total of 146 s over 2 flights.

The PA final design is a high wing monoplane with an empty weight of 0.947 kg and a maximum flight speed of 20 m s−1 .

Manufacturing materials mainly involved carbon fiber, balsa and foam. The aircraft is estimated to be able to complete 3 laps in 132 s for Mission 3. The PA can split into two sub-assemblies, the wing / landing gear group and the propeller / fuselage / boom / empennage group. The team determined that the PA can be assembled and payload loaded within the 2 minute requirement for the Bonus Mission.

The estimated RAC for the design is 0.872.

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2

Management Summary

The University of Glasgow Design, Build, Fly team has members from first year undergraduate to postgraduate level students, of which over a third are either freshmen, sophomores or juniors. Besides the competition requirements, this composition is intended to sustain the project in future years as experience gets passed on to the younger members.

Overall, the team has 23 members from 12 countries, supervised by faculty advisers. The project is an extra-curricular activity since students have to work on it besides their usual curricula and academic courses. Being the second entry for the University of Glasgow, the management team decided to expand the team’s capabilities by acquiring additional equipment and tools, introducing more advanced manufacturing and testing techniques, as well as seeking sponsors to obtain the required funding. The number of engineering team members this year was limited to 20 based on previous year’s experience. The number of applicants significantly increased from last year, making the application process rather demanding and competitive. The organizational structure of the team is shown in Figure 1.

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

Team meetings were scheduled every week and Slack, a cloud based messaging tool, was used along with Google Drive to facilitate information sharing and off campus communications. The Project Manager was responsible for planning and following up on the master schedule and overall project goals. Moreover, the project manager took final responsibility for all major decisions and supervised the budgeting process. The Operations Manager assisted the Project Manager, primarily supervised the projects technicalities, maintained internal communications, and coordinated the team’s resources according to the master schedule. The Chief Engineer reported to the Project and Operations Managers to provide an extensive checking system for all critical engineering decisions. The team’s advisers are the former project and operations managers from the University of Glasgow DBF 2014/2015 team entry, and their main task was to provide guidance and share their gained experience, in addition to monitoring the design approach and decisions. The Treasurer and Logistics Manager maintained record of all the project’s assets, income and expenditure, issued purchase orders, and managed inventory.

The sponsorship team, consisting of two penultimate year business students, worked independently from the rest of the team and took the responsibility of maintaining relations with commercial sponsors. This was especially important as the team sought establishing its reputation amongst potential industrial and commercial sponsors.

Furthermore, the team was divided into four design sub-teams: Aerodynamics, Propulsion, Mechatronics and Structures.

Each sub-team had a designated leader, who would manage the sub-team’s activities, ultimately reporting to the chief engineer, who relays the information to the project and operations managers. The outline of sub-team tasks was as follows: Aerodynamics Sizing of aircraft components, performance and stability estimates

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Propulsion Motor, ESC and Propeller sizing, testing and selection Mechatronics Aircraft battery, servos, receiver and transmitter sizing, testing and selection Structures Design and materials for aircraft fuselage, landing gear, empennage and supervising aircraft assembly

2.1

Milestones Chart

Since time is one of the factors that affect the design process, a general schedule for the project was setup to organize the necessary deadlines. The schedule includes everything from early design stages to flight testing. The project’s gantt chart is shown in Figure 2.

Figure 2: Master schedule for the project.

Discrepancies between the planned and actual schedule on some occasions were caused primarily by extension of preliminary design period due to the change of aircraft concepts from a flying wing configuration to a conventional configuration. The team has been accumulating lessons learned style information to preserve experience that will be used in forthcoming competitions.

3

Conceptual Design

This section describes the initial stage of designing the two aircraft. The goal is to identify the optimal configuration that would provide the highest mission score and the lowest RAC.

3.1

Design Constraints

From the design requirements, the following main design constraints have been identified:

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1. Only Nickel Metal hydride (NiMH) or Nickel Cadmium (NiCd) batteries are allowed, although there is no limit on a battery mass 2. Both aircraft have to be able to takeoff within 100 ft (30 m) 3. Both aircraft should be able to land safely without taking significant damage 4. The two aircraft must be capable of performing all the missions (a) The MSA must be able to fly without payload (b) The MSA must be able to internally carry the PA as one or more sub-assemblies (c) The PA must be able to internally carry a 32 oz. Gatorade bottle (d) All payload must be adequately secured 5. Both aircraft should withstand a wingtip test when fully loaded The competition course layout with turn zone and runway is shown in Figure 3. Total length of the course including turns is approximately 3000 ft (1 km).

360˚ Turn

Cruise 1000ft (300m) 180˚ Turn

Runway 100ft (30m)

Not to scale

3.2

Figure 3: Layout of the competition course. Not to scale.

Scoring Outline

In order to understand the scoring sensitivity, the scoring formulae and the missions associated with them need to be studied.

The Total Score is given by Total Score =

3.2.1

Written Report Score × Total Mission Score RAC

(1)

Rated Aircraft Cost

Reducing RAC is one of the means of achieving a higher score, as it divides the product of Report Score and Total Mission score. RAC can be calculated using the formula: RAC = EW1 × WtBattery 1 × Ncomp + EW2 × WtBattery 2

(2)

where EW1 is the weight of the PA ready to fly but without payload, WtBattery 1 is the battery weight for the PA, Ncomp is the number of sub-assemblies the PA is broken into for the delivery flight(s), EW2 is the weight of the MSA ready to fly but without a payload and WtBattery 2 is the battery weight for the MSA.

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3.2.2

Total Mission Score

Total mission score is given by Total Mission Score = MF1 Ă— MF2 Ă— PF + Bonus

(3)

where MF1 and MF2 are scores for the MSA flight missions, PF is the score for PA flight mission and Bonus is the bonus task score for PA ground mission. All flight missions require takeoff within the prescribed field length and a successful landing.

The components of total mission score are as follows:

MSA Arrival Flight, MF1

The first mission requires the MSA to takeoff, fly 3 laps without a payload around the competition

course within 5 minutes and land. The possible scores for the mission are either MF1 = 2.0 for completing the mission else a score of MF1 = 0.1 is awarded.

MSA Delivery Flight, MF2

The second mission requires the MSA to carry the PA sub-assemblies around the course. One

or more parts of the PA can be carried in a single flight. The MSA has to takeoff, fly one lap and land for each sub-assembly delivery flight. After landing, the MSA has to taxi to a payload loading area where the installed sub-assembly is removed and a new sub-assembly is loaded, if required. This is repeated until all the PA sub-assemblies have been delivered. All the PA parts need to be carried internally and the mission has to be completed within 10 minutes. The score awarded for the mission is MF2 = 4.0, if all sub-assembly delivery flights are successfully completed or MF2 = 1.0 if at least one sub-assembly group is successfully delivered. If no flights have been attempted or completed successfully MF2 = 0.1 is awarded. If a partial score is received (MF2 = 1.0), the aircraft may attempt a single retry of the mission.

PA Flight, PF For the third mission, the PA has to carry a single 32 oz Gatorade bottle of any flavor for 3 laps around the course within a time limit of 5 minutes. The payload must be carried internally and secured sufficiently. Approximate specifications of the payload are: height 8.2 in (20.8 cm), maximum diameter 3.7 in (9.4 cm) and mass 2 lb 3.9 oz (1020 g). The payload can be seen in Figure 4. A score of PF = 2.0 will be awarded for completing all mission objectives. If the PA flies

Figure 4: 32 oz Gatorade bottle used as the payload for mission 3.

less than the required number of laps or exceeds the time limit, a score of PF = 1.0 is awarded. In the event of no flight attempts being made or failure to successfully complete a flight attempt, PF = 0.1 is awarded. If a partial score is received (PF = 1.0), the aircraft may attempt a single retry of the mission.

Bonus Mission

The bonus mission requires the team to assemble the PA from its sub-assemblies within 2 minutes; this

includes installing and securing the payload. The bonus mission can only be attempted if the MSA successfully completes Mission 2 (MF2 = 4.0). The aircraft will receive a score of Bonus = 2.0 if the PA is assembled within the specified time limit and passes the wing tip test as well as the control systems check; otherwise Bonus = 0.0.

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3.3

Total Score and Sensitivity

From Section 3.2, it can be concluded that if all mission requirements are met, the only opportunity to improve the score lies with the Written Report Score and RAC. Therefore, a sensitivity analysis has been performed to assess the impact of different configurations on the RAC. The RAC component of the PA is primarily defined by the payload it has to carry and Ncomp , so the only way to improve it is by having a well optimized configuration, which is the goal of the preliminary and detailed design stages. The RAC component of the MSA, on the other hand, is directly linked to how MF2 is approached.

Three options for MF2 were analyzed as the most viable and are presented in Table 1.

Option

Ncomp

MF2 Number of flights

I

1

1

II

2

2

III

2

1

Table 1: Options considered for MF2.

In general, Option I will result in having a larger and heavier MSA, since the entire PA has to be carried internally, however the RAC will benefit from Ncomp = 1. Option II will have a lighter MSA since only a half of the PA has to be carried per flight, this also results in having a lighter battery. This is due to the fact that a smaller payload would require a smaller MSA, bearing in mind that this is a general assumption that is used to conduct the analysis. Although this option still incurs the penalty of increasing Ncomp to 2. Option III allows for improved overall efficiency of the MSA, as 2 separate components could be packed more densely inside the payload compartment, however the penalty incurred by Ncomp = 2 remains. An option with 3 components was also considered for different numbers of laps in MF2 , however this always produced a higher RAC, since with Ncomp = 3 in EW1 × WtBattery 1 × Ncomp outweighs any benefit obtained by reducing EW2 × WtBattery 2 , therefore this was not described in detail.

Consequently, Options I and II were the first to be evaluated with the analysis primarily based on the ratios of EW to Wtpayload and EW to Wtbattery . Those ratios were estimated based on experience and previous reports for the 3 options and were determined to be the main variables influencing the RAC. The base assumptions can be found in Table 2. Note, it was assumed that it would be possible to split the PA for Options II and III into 2 components of approximately equal masses and dimensions.

EW1

1.02

PA mass

Wtbattery1 /EW1

30%

Baseline battery ratio

Wtbattery1

0.31

PA battery mass

WtPayload, I

0.71

Option I payload mass

WtPayload, II

0.36

Option II payload mass (per flight)

WtPayload, III

0.71

Option III payload mass

Table 2: Experience based initial assumptions.

With these assumptions, it was possible to calculate the RAC for Options I and II as a function of

EW2 Wtpayload .

Figure 5 shows the

relationship between payload ratios for Option I and II the RAC.

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1.8

1.6

1.4

RAC

1.2

Option II wins

1

0.8

Option I 0.6

Option II

Option I wins 0.4 1

1.2

1.4

1.6

1.8

2

EW2/ Wt

2.2

2.4

2.6

2.8

3

payload

Figure 5: RAC vs

EW2 .

Wtpayload

This shows that as long as the MSA for Option II has a payload ratio greater than 1.65, it is worth going forward with Option II. For instance, given that

EW2 Wtpayload

= 2.3, the structure of the MSA for Option I would have to be 20% lighter in order to

match the RAC of Option II. Considering that having to carry the whole PA internally would require larger dimensions of the MSA, it is safe to assume that it would be practically impossible to achieve a better payload ratio than for Option II. To investigate whether it was possible to achieve a ratio lower than 1.65 for Option I, preliminary flying wing PA and conventional MSA design analyses were produced. It was confirmed that it would be impossible to achieve the aforementioned ratio, or better, based on team’s expertise. It should also be noted that the RAC curve for Option I ratio is steeper, resulting in higher sensitivity to design imperfections, as opposed to Option II which can be considered as a generally safer design solution.

Next, Option II was compared to Option III. Since Mission 1 requires the MSA to be able to fly 3 laps, it was first assumed that meeting this requirement would automatically allow flying 2 laps in MF2 for Option II. However, the concern was that NiMH produce less power after some discharge, resulting in a requirement for a higher capacity battery in order to produce the required power for the second takeoff. This would introduce a battery mass penalty. In addition to that, the Option III MSA could achieve a better payload ratio, as the volume required to fit one part of the PA would also allow fitting a second part at the same time for various configurations. Those two factors combined could potentially make Option III more advantageous.

As well as

EW2 Wtpayload ,

battery mass ratio

Wtbattery 2 EW2

was included in the analysis to select the optimal configuration. The RACs

for Options II and III were then calculated to obtain baseline estimates, where

EW2 Wtpayload

ratios were influenced by experience

gained in previous years. From that baseline, the battery mass fraction of Option II was changed to match the RAC of Option III. The headroom for increases in battery mass for Option II can be found in Table 3.

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Possible payload mass ratios for II and III

Baseline battery mass fraction

Adjusted for maximum battery mass fraction

WtBattery 2 /EW2II

30%

30%

30%

30%

30%

WtBattery 2 /EW2III

30%

30%

30%

30%

30%

EW2 / WtpayloadII

3

4

5

4

5

EW2 / WtpayloadIII

2

3

4

2.5

3

RACII

0.97

1.24

1.58

1.24

1.58

RACIII

1.24

2.00

3.07

1.58

2.00

WtBattery 2 /EW2II

54%

67%

77%

47%

43%

WtBattery 2 /EW2III

30%

30%

30%

30%

30%

EW2 / WtpayloadII

3

4

5

4

5

EW2 / WtpayloadIII

2

3

4

2.5

3

RACII

1.24

2.00

3.07

1.58

2.00

RACIII

1.24

2.00

3.07

1.58

2.00

Battery mass increase headroom

79%

125%

156%

56%

44%

Table 3: Battery sensitivity analysis for Options II and III.

As it can be seen, there is sufficient battery mass penalty headroom for even the most conservative payload ratios, which implies that Option II will allow the highest RAC to be achieved, and thus the highest overall score possible.

Mission requirements in general are translated to aircraft requirements in Table 4.

Mission Requirement

Aircraft Requirement

100 ft takeoff

Sufficient lift and thrust-to-weight

Fly 3 laps empty and deliver the PA sub-assemblies

Battery capacity

Fly 3 laps with the PA and payload

Battery capacity

Minimize EW

Optimize the design, use efficient materials

Minimize RAC

Split the PA in 2 sub-assemblies as efficiently as possible, aggressively optimize the battery performance

Fast loading and unloading of the MSA in M2

Easy access to payload compartment

Assembly of the PA below 2 minutes

Easy mounts and fixes

Table 4: Mission requirements translated to aircraft requirements.

3.4

Configuration Selection

To reach the final conceptual design, the team considered various aircraft configurations. The main configurations considered are shown in Table 5. Many of the combinations were quickly deemed to be impractical and only the most viable overall configurations are considered below.

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Component

Configurations

Main Criteria

Wing Layout

Flying wing

Monoplane

Wing Attachment

Low

High

Tail Type

T-tail

V-tail

Number of Motors

1

2

Motor Location

Tractor

Pusher

Both

Dimensions, CG

Landing Gear

Tricycle

Taildragger

Bicycle

Mass, ability to split PA

Biplane

Blended

Mass, manufacturing, Ncomp Stability

Conventional

H-tail

Mass, dimensions Mass

Table 5: The configurations considered for the two aircraft at a conceptual design level. Bold indicates selected configuration and the main criteria column lists the main reason(s) for the selection.

Mission 3 requires the PA to carry the specified payload internally. Thus the team decided to design the fuselage to follow the dimensions of the payload as closely as possible to reduce the mass and size of the aircraft. On the other hand, the main purpose of the MSA is to carry the sub-assemblies of PA internally around the competition course in Mission 2. Hence the driving factor in the MSA design is to fit the PA sub-assemblies internally with minimal structural requirements in terms of mass and dimensions. In turn this requires the PA to be as light and compact as possible while still being able to carry out Mission 3.

From sensitivity analysis, it was concluded that two flights in Mission 2, carrying PA in two separate sub-assemblies would provide the highest score based on team’s knowledge and understanding. Therefore to minimise the mass of the MSA, the two PA sub-assemblies should be of similar mass and dimensions. This approach avoids wasting internal volume of MSA and ensures that the MSA payload requirement is as low as possible. The heavier piece would be carried first to take advantage of a fresh battery pack and make the second takeoff with partially discharged batteries less power demanding.

A flying wing design was briefly investigated initially to determine if a one flight, one component approach could be taken for Mission 2. However, it was concluded that it was not realistic to obtain high enough structural efficiency to justify this option. Furthermore an aircraft with conventional configuration can be designed, with relative ease, to split into two sub-assemblies of similar dimensions and masses by having wing / landing gear as one group and the propeller / fuselage / boom / empennage as the other group, diminishing the advantage of a flying wing configuration. Also conventional design was judged to be a less risky option. Therefore, for the PA, the wing span should closely align with the overall length of the aircraft and the wing chord and landing gear height should be close to the diameter of the PA fuselage. The same applies for the horizontal and vertical tails. The layout is presented in Figure 6, showing how the sub-assemblies will be loaded inside the MSA fuselage. The view is idealized and shows two of the main constraints of fitting the PA sub-assemblies into the MSA while maintaining the same bounding box.

The approach chosen first drove the team to adopt a tail dragger landing gear for the PA since this would position the majority of the landing gear mass to the main gear that would be easy to attach to the wing. Also, a tail dragger design is generally the lightest landing gear option. No other landing gear arrangement would allow majority of landing gear mass to be attached to the wing. For example, tricycle landing gear requires a nose gear that would have to be located far in front of the wing, and would be much larger than the tail wheel of a tail dragger configuration for a given aircraft.

A high wing configuration was chosen to give increased stability by having the aircraft center of gravity (CG) below the wing and to make sure that the tail assembly would be subject to less aerodynamic shadowing, as a conventional tail was adopted. Shadowing was considered important because the aspect ratios for vertical and horizontal tails would be very low in order to make sure they do not extend far outside the PA fuselage diameter. This immediately made using a T-tail a very inefficient option since the horizontal tail should be high enough on top of the vertical tail to avoid deep stall, where the fuselage and wing shadow both tails completely. V-tail has the possibility to reduce drag by eliminating structural junctions that cause aerodynamic interference. However the strong coupling of lateral and longitudinal stability was to be avoided with U NIVERSITY OF G LASGOW

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PA wing and landing gear

Wing Chord

Vertical tail

Fuselage

Chord Fuselage

Span MSA fuselage outline

Figure 6: Conceptual arrangement of the propeller / fuselage / boom / empennage group and the wing / landing gear group. The two views show the main constrains of fitting PA sub-assemblies into MSA.

the PA since the aircraft would already have complicated stability characteristics due to the dimensional limitations of the empennage that has to fit into the MSA payload bay. Also, in practice, it is difficult to optimise a V-tail such that significant mass savings or aerodynamic improvements can be made. Furthermore, a conventional tail is the lightest option of the three.

For propulsion of the PA it was decided to limit the propeller diameter in order to maintain the same bounding box around the PA. Tractor configuration was determined to be the most efficient option since with pusher configuration it would be difficult to balance the CG and it would cause problems with the roll angle, as tail dragger gear was selected.

The main challenge of the MSA design is to fit the two PA sub-assemblies internally while keeping the mass as low as possible. To achieve minimum mass, a tail dragger configuration was also selected for the MSA. Although tricycle gear provides better ground handling qualities, this was not thought to be a major disadvantage this year considering the number of servos is not limited. A high wing design was chosen due to generally having better stability characteristics. Also, the top surface of the wing, which produces the majority of the lift, can be left undisturbed leading to more efficient lift generation.

Less benefit from ground effect in comparison to a low wing configuration was not considered to be large enough to outweigh the benefits of a high wing. Loading of the sub-assemblies was decided to be done from the back of the plane to minimize the opening in the fuselage. Since the loading procedure needs to be reasonably fast so that Mission 2 can be completed in time, this approach ensured that high wing would not cause problems.

To make balancing the MSA easier, tractor configuration was selected for propulsion. Tractor configuration allows higher propulsive efficiency because the propeller receives clean airflow in front of the aircraft, whereas pusher configuration would suffer from the fuselage and wings disturbing the flow. Also, propeller slipstream provides high momentum flow over the wing which improves lift, especially during takeoff. From previous experience, there was no significant benefit observed from having more than one motor. The main reason for this is an increase in mass from multiple motors, propellers, ESCs and requiring more wiring. Thus, only one motor was to be used.

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4

Preliminary Design

4.1

Design Methodology

The preliminary configuration parameters of both aircraft were obtained through using a constraint analysis approach from Gudmundsson.1 This approach uses analytical equations which express aircraft power loading wing loading,

W S

P W

, as a function of aircraft

, and other important parameters such as minimum drag coefficient CD0 , induced drag constant k, propulsive

efficiency ηp , climb rate VV , climb velocity Vcmb , turn velocity Vtrn , turn load factor n and maximum lift coefficient CLmax . In addition to these parameters the aircraft power loading is also affected by other various constraints including laps that have to be flown, payload, and safety factors. The constraint analysis approach described above was used for both aircraft. As there is a requirement that one aircraft should carry the other split in components, the constraint analysis of both aircraft influence each other.

4.1.1

Airfoil Selection

Airfoil selection is one of the most crucial parts during the preliminary design of an aircraft since it has a huge impact on several important parameters. It was decided that no high lift devices will be used which resulted in a requirement for the airfoil to have high maximum lift coefficient. A high maximum lift coefficient is favorable for the take off performance of both the PA and MSA due to the fact that it ”lowers” the take off constraint curve present in the constraint analysis figures of the PA and MSA, shown in Figures 11 and 12 respectively. This in turn lowers the required power loading for both aircraft which results in reduction of their empty mass.

A second important requirement was that the airfoil should be suitable for low Reynolds number flow, usually less than 500,000 and that it should exhibit trailing edge stall. Such flow aerodynamics, require great attention for a number of reasons. Firstly, operating at the aforementioned flight regimes is often associated with hysteresis phenomena. Hysteresis represents the current flow state dependency on previous time states and is particularly important for stall recoveries, spin flight or high gust conditions. This unsteady phenomenon highly increases the drag and decreases the lift when the angle of attack of an airfoil is reduced in post stall operations, possibly resulting in loss of control. Furthermore, hysteresis is strongly coupled with the second phenomenon of low Reynolds number airfoils - laminar separation bubbles. Laminar separation bubbles form when airflow that is laminar separates because of high adverse pressure gradients and when separated, transitions to turbulent flow. It then curves back and reattaches to the surface, creating a shallow region of reversed flow.

Such bubbles increase drag, which can become a few magnitudes bigger than the drag of the airfoil without the bubble, and are to be avoided. Additionally, such bubbles are responsible for leading edge stall resulting in abrupt changes in the lift coefficient near the stall region.

The third important requirement was that the airfoil should have simple geometry so that it can be manufactured precisely.

After analyzing a list of suitable airfoils four options were chosen. NACA4412, Clark Y, NACA 0012 and Eppler 171. The first two airfoils are very common in model aircraft, furthermore they have a relatively high lift coefficient, favourable drag coefficient and are very easy to manufacture. The selected airfoils and their aerodynamic characteristics are shown in Figures 7 and 8 respectively.

1 S.

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

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0.1 0.05 0 NACA 4412 Clark Y NACA 0012 Eppler 171

-0.05 -0.1 -0.15 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Figure 7: Geometry of researched airfoils.

1.5 100

Eppler 171 Clark Y NACA 0012 NACA 4412

1

80 60

0.5

Cl

Cl/Cd

40 20 0

-0.5 Eppler 171 Clark Y NACA 0012 NACA 4412

-20 -40 -60 -10

0

-1

-1.5 -5

(a)

0

5

Angle of attack (deg)

Cl Cd

10

0

15

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

Cd

(b) Cl vs Cd

vs Angle of Attack

0.04

1.5

0.02 1

0

Eppler 171 Clark Y NACA 0012 NACA 4412

0.5

Cl

Cm

-0.02 -0.04

0

-0.06 -0.5

-0.08 Eppler 171 Clark Y NACA 0012 NACA 4412

-0.1 -0.12 -10

-5

0

5

10

Angle of attack (deg)

(c) Cm vs Angle of Attack

-1

15

-1.5 -10

-5

0

5

10

15

Angle of attack (deg)

(d) Cl vs Angle of Attack

Figure 8: Aerodynamic characteristics of researched airfoils.

4.2 4.2.1

Parameter Selection and Estimation Drag Estimation

Zero-lift drag buildup model from Roskam2 and Gudmundsson3 was used to determine drag, component by component, using 20 m s−1 as a reference airspeed. First, wetted areas of the exposed components were calculated and the fuselages of both aircraft were approximated as boxes that are slightly bigger than the payload. The skin friction coefficients were then estimated, depending on the component location and the Reynolds number of the flow over the component. The drag of the 2 J. Roskam. Airplane Design: Part VI: Layout Design of Cockpit, Fuselage, Wing and Empennage : Cutaways and Inboard Profiles. Airplane Design.

DARcorporation, 2002.

3 Gudmundsson, General Aviation Aircraft Design: Applied Methods and Procedures.

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landing gear was estimated with a method described by Gudmundsson.4 The estimated drag coefficients for PA and MSA components are shown in Tables 6 and 7, shown below.

Component

Wetted area (m2 )

Form factor

Equivalent flat plate area (m2 )

Parasitic drag coefficient

Wing

0.4332

1.4556

0.003704

0.0183

Horizontal tail

0.0831

1.3447

0.000708

0.0035

Vertical tail

0.0299

1.3447

0.000263

0.0013

Fuselage

0.0671

4.8462

0.001741

0.0086

Landing gear

-

-

0.000607

0.0030

Total

-

-

0.007023

0.0347

Total including factor of 1.06

-

-

0.007444

0.0368

Table 6: PA drag buildup.

Component

Wetted area (m2 )

Form factor

Equivalent flat plate area (m2 )

Parasitic drag coefficient

Wing

0.5870

1.4556

0.063470

0.0184

Horizontal tail

0.1449

1.3447

0.001363

0.0047

Vertical tail

0.0759

1.3447

0.000667

0.0023

Fuselage

0.9239

1.3945

0.005479

0.0189

Landing gear

-

-

0.001218

0.0042

Total

-

-

0.072197

0.0485

Total including factor of 1.06

-

-

0.076529

0.0514

Table 7: MSA drag buildup.

Using the XFLR5 software, an average Oswalds efficiency factor of eP A = 0.8950 for the PA and eM SA = 0.9630 for the MSA was determined and drag polars for both were obtained. They are shown in Figures 9a and 9b.

1.2

1.2 Mission 2 CL (Heaviest component) Mission 1 CL

1

0.8

0.8

0.6

0.6

CL

CL

Mission 3 CL

1

0.4

0.4

0.2

0.2

0

0

0.01

0.02

0.03

0.04

0.05

0.06

0.07

0.08

0.03

0.04

CD

0.05

0.06

0.07

0.08

CD

(a) PA CL versus CD

(b) MSA CL versus CD

Figure 9: Drag polar for PA and MSA.

4 Gudmundsson,

General Aviation Aircraft Design: Applied Methods and Procedures.

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4.2.2

Mass Estimation

During the preliminary design stage both the wing and motor mass vary due to changes in wing and power loadings. To account for these changes, wing density per area was first estimated based on previous reports, describing similar aircraft configurations with similar aspect and taper ratios. This value was later refined after the weight of the PA and MSA wing was estimated from their CAD models and first prototypes. As for motor mass, a variety of motors were researched and their mass and power recorded. From these values a motor mass versus power figure was generated and a linear relationship obtained.

Mass values for other aircraft components were estimated in a similar manner - existing components were analyzed and their masses recorded after which an average value was calculated.

Similarly, battery mass also changes at different power loadings. This is due to the fact that at larger power loadings more powerful motors are required hence more powerful batteries. The battery mass estimation, however, was more difficult than the estimations for the mass of the wing and the motor as it is dependent on the aircraft performance. The aircraft performance was analyzed during different sections of the mission - climb, turn, and cruise. From this analysis the peak power requirement was estimated. Based on this peak power value and battery cell specifications, a relationship was developed which estimates the battery mass from the peak power, battery C rating and energy density.

Battery Mass =

Maximum Power Requirement Battery C rating × Battery energy density

(4)

The mass estimations evaluated with the previously mentioned methods for both the PA and MSA are outlined in Tables 8 and 9. The aforementioned methods were used to estimate the mass of both aircraft at various power loadings.

Component

Mass (kg)

Servo

0.03

Receiver + Receiver battery + Fuse

0.06

ESC

0.05

Propeller Motor Wing Battery Fuselage Landing gear

0.04 P ×W Aircraft mass× 3140.2400 Aircraft mass× 0.9874 W ×S Maximum power required 880

0.2 0.05

Table 8: PA mass estimations. Note that these values were obtained from the methods discussed previously.

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Component

Mass (kg)

Servo

0.03

Receiver + Receiver battery + Fuse

0.06

ESC

0.05

Propeller Motor Wing Battery

0.04 P ×W Aircraft mass× 3140.24 Aircraft mass× 0.9874 W ×S Maximum power required 880

Fuselage Landing gear

0.5 0.07

Table 9: MSA mass estimations. Note that these values were obtained from the methods discussed previously.

4.3

Constraint Analysis

The most important mission constraints for both the PA and MSA include: • Empty mass and battery mass • Takeoff distance • Turn speed • Cruise speed • Climb speed • Rate of climb For the MSA, an additional crucial mission constraints are payload dimensions and weight. Similar constraints apply for the PA. All of the above constraints are considered by the constraint analysis algorithm which determines the minimum aircraft mass at a particular

P W

and

W S

for which the constraints are satisfied.

Constraint Analysis Algorithm As discussed earlier the constraint analysis algorithm takes into account various constraints which are imposed by the missions. Therefore, each aircraft must satisfy a different set of constraints. By taking into account each constraint the algorithm iteratively calculates the power required for takeoff, turn, and climb at a particular wing loading.

As the wing loading changes, other dependent parameters change accordingly. Such parameters include areas and masses of the wing and the horizontal and vertical tail, as well as motor and battery mass. Accordingly, these changes influence the aircraft’s minimum drag coefficient, induced drag constant, all of which have an effect on the equations used in the constraint analysis. As a result these equations do not only depend upon wing loading, but also on other parameters which are influenced by wing loading. Consequently, an iterative approach is used by the algorithm in order to account for the previously mentioned dependencies. For each power and wing loading the algorithm iteratively changes the above mentioned parameters until the estimated empty mass of the aircraft converges. The result yielded by the algorithm is a family of different curves for takeoff, turn and climb power loadings which are plotted against the estimated empty mass of the aircraft. A block diagram of the algorithm is shown below in Figure 10.

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Geometry Wing geometry, horizontal and vertical tail volume coefficients and geometry, distance between aerodynamic centers of wing and horiozntal and vertical tail

Mass Typical components masses, wing, motor, and battery densities

Design requirements ROC Cruise airspeed Climb airspeed Turn airspeed Take off distance Maximum achievable lift coefficient

Preliminary mass estimation generate mass matrix

Generate power loading and wing loading matrix

Preliminary drag buildup generation of drag matrix

Safety factors Battery mass Take off distance

Power loading (P/W), wing loading (W//S)

Mass=Estimated mass (preliminary) Mass (previous) = 0

YES

Inner loop

Mass - Mass (previous) <= 0.01 NO Mass (previous) = Mass

Recalculate wing, horizontal tail, vertical tail and motor masses due to change in mass using P/W and W/S and respecive densities

Analyse take off, climb, turn and maximum airspeed performance

Mass

Recalculate wing geometry, horizontal and vertical tail geometry due to the change in mass using W/S

Find required power

Recalculate drag due to change in wing, horizontal and vertical tail

Calculate battery mass

Outer Loop (Loops through the power, wing loading, mass and drag matrix) Minimum mass configuration satisfying all design requirements

P/W, W/S, wing, horiozntal tail, vertical tail geometry, motor power requirement

Figure 10: Block diagram of constraint analysis algorithm.

4.4 4.4.1

Mission Constraints Production Aircraft

The preliminary design of the production aircraft was carried out first due to the fact that the aircraft, when split in components, is the payload of the MSA. This imposed a constraint on the production aircraft to be as small and light as possible. However making the aircraft smaller has an effect on other constraints which are discussed below.

Takeoff

In case of the PA, takeoff is the most critical constraint. It dictates the minimum wing size with the selected

maximum lift coefficient. The bigger the CLmax , the lower the stall speed and wing area, which would ultimately decrease the total mass.

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Turn

The aircraft is designed to fly at a design speed of 14 m s−1 during the turns. This speed is less than the maximum

speed of 20 m s−1 during the straight sections of the course. Having a lower turning speed reduces the load factor which in turn allows reduction of wing strength and consequently mass.

Climb

The PA was designed to be able to achieve a rate of climb of 2 m s−1 at an airspeed of 16 m s−1 .

Constraint Analysis results

2

1.4

60 40

0.9

20

9 1.

8 1.

5

1.

4

1.

120

2

1.1

80 60 40 20

0

1.2

100 1.

1.1

1.2

140

7 1.

6 .

1

1.1

1.2

160

1.7

1.4

1.5 1.6

1.6 1.5

180

0.9

1.4

80

1.5

1.1

100

1.7

1.8

120

1.6

1.4

1.6 1.7

140

2

160

1.5

P/W - Power loading (W/kg)

180

Empty weight Takeoff (24.38 m) 1.6 Climb (16.00 m/s ROC 2m/s) Turn (14.00 m/s) 1.5 Maximum speed (20.00 m/s) 1.4 Best range Minimum mass point Selected mass point

200

1.9 8 1.

P/W - Power loading (W/kg)

200

0 4

6

8

10

12 2

W/S - Wing loading (kg/m )

5

10

15

20

25

30

35 3

W/S - Wing cubic loading (kg/m )

Figure 11: PA constraint analysis. Note that the mass values below the constraints have no physical significance. This is due to algorithm showing that the aircraft stalls when power available is less than the one for a specific constraint. Additionally, the mass lines are discrete, because of battery mass increasing not linearly, but as an integer multiplier of battery cell mass.

Safety factors and estimations

Value

Takeoff distance safety factor

0.8

Battery mass safety factor

1.2

Estimated propulsive efficiency

0.6

Estimated ground friction coefficient

0.025

Maximum achievable lift coefficient at reference airspeed (NACA4412)

1.5

Table 10: Safety factors and estimations for PA used in the constraint analysis.

4.4.2

Manufacturing Support Aircraft

Similar design parameters were selected for the MSA as it was determined that they are the optimum for the flight course.

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1.9

2

1.8

1.9

1.7 1.6

1.8 1.7

80

1.5

6

60

1.

1.4

3.54

3

1.6

1.7 60

1.5

1.4

20

0

1.5

1.5

1.9

20

1.7

100 80

1.8

2

40

2

40

120

2

100

3

120

140

2.5

140

160

1.6

2

1. 1.9 8

5

2.

P/W - Power loading (W/kg)

160

2.5 Empty weight Takeoff (24.38 m) Climb (16.00 m/s ROC 2m/s) Turn (14.00 m/s) 2 Maximum speed (20.00 m/s) Best range 1.9 Minimum mass point

2.5

4 3.5

180

2.5

P/W - Power loading (W/kg)

180

200

2.5

3

3

200

0 4

6

8

10

12

W/S - Wing loading (kg/m 2 )

5

10

15

20

25

30

W/S - Wing cubic loading (kg/m 3 )

Figure 12: MSA constraint analysis. Note that the mass values below the constraints have no physical significance. This is due to algorithm showing that the aircraft stalls when power available is less than the one for a specific constraint. Additionally, the mass lines are discrete, because of battery mass increasing not linearly, but as an integer multiplier of battery cell mass.

Safety factors and estimations

Value

Takeoff distance safety factor

0.8

Battery mass safety factor

1.4

Estimated propulsive efficiency

0.6

Estimated ground friction coefficient Maximum achievable lift coefficient at reference airspeed (NACA4412)

0.03 1.5

Table 11: Safety factors for MSA used in the constraint analysis.

4.5

Trade-off Studies

In the constraint analysis Figures 11 and 12 shown earlier, the minimum empty mass point which satisfies all constraints can be seen. For this point based on the input parameters provided to the algorithm, shown in Tables 14 and 16, output parameters, shown in Tables 15 and 17, such as wing area, horizontal tail area, empty mass, and others were calculated.

However the output parameters at the minimum mass point for the PA, though optimal, do not satisfy the previously imposed condition that this aircraft should be as small as possible in order to easily fit, when broken down in components inside the MSA. Due to this reason, the parameters at the minimum mass point of the PA were not used but instead, parameters at a desired wing loading and power loading were selected. The new parameters ensured that the wing span and the horizontal tail span are small enough, so that the PA can fit into the fuselage of the MSA. The reduction of the wing and horizontal tail span was achieved through increase in wing and power loading. In Figure 11, showing the constraint analysis results for the PA, marked with a circle, is the selected point for the desired wing and power loading. The drawback to reducing the span of the wing was that power had to be increased in order that the takeoff constraint could be met. This issue arises due to the fact that wings with higher wing loading stall at higher speeds. For this reason different values of wing loadings and their effects are discussed below.

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High Wing Loading

Low Wing Loading

Stall Speed

High

Low

Takeoff Distance

Long

Short

Maximum Lift to Drag Ratio

High

Low

Gust Response

Good

Bad

Low

High

Weight

Table 12: Effect of different wing loadings on aircraft.

Loadings up to 3 kg m−2 are for gentle flying, 3 kg m−2 to 6 kg m−2 for trainers and above 6 kg m−2 for aerobatics. However, Reynolds5 suggests that wing cubic loading factor, defined as weight divided by the square root of the wing area, would be more appropriate to estimate the handling qualities for RC airplanes. Including the linear term, dividing by root of area, i.e. some reference length, in the wing loading equation introduces the airplane wing area as another factor. This makes it easier easier to categorize airplanes according to their ease of control.

Wing cubic loading for the PA at the minimum mass point is 16.33 kg m−3 and at the selected wing and power loading point is 22.44 kg m−3 . According to Myers6 the aircraft would categorize as an expert sport for the former and as an expert only sport for the latter. It was decided that as long as the pilot has experience flying aircraft, handling qualities similar to an expert only sport aircraft are manageable, hence the desired wing and power loading point was selected. The selection of this point also satisfied the constraint of the PA to be small.

The MSA has a wing cubic loading at the minimum mass point which varies from 9.99 kg m−3 when it is empty to 14.30 kg m−3 with full payload. Again according to Myers7 the aircraft would categorise as a sport type when empty and as an advanced sport when flying with payload. As the the type of the production aircraft was selected to be expert only, the minimum mass point for the MSA was selected.

Average wing cubic load factor (kg m−3 )

Level

Description

1

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

2.39

2

Includes mostly backyard type models that can be flown indoors in larger venues and outside in low wind conditions, includes a few internal combustion powered planes

4.1

3

Includes park flyers, sailplanes, biplanes, 3D planes

5.99

4

Includes sport types, biplanes, scale, a few 3D planes, pattern, largest level

8.51

5

Includes advanced sport types, sport scale and sport scale warbirds, some twins

11.25

6

Includes expert sport types, scale, scale warbirds, twins

14.31

7

Includes planes for the expert flier only, twins and multi-motor, true scale, warbird Table 13: RC aircraft general handling quality levels and respective wing cubic loadings.

5 Francis 6 Ken

Reynolds. Wing Cube Loading. Model Builder. 1989.

Myers. Club Newsletter. 2014. URL: http://www.theampeer.org/M1-outrunners/M1-outrunners.htm#CWL.

7 Ibid.

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4.6

Estimated Parameters Important input parameters Maximum lift coefficient

1.5

Propulsive efficiency

0.6

Wing aspect ratio

5.5

Horizontal tail aspect ratio

1.6 1

Vertical tail aspect ratio Vertical tail taper ratio

0.5

Horizontal tail volume coefficient

0.6 0.04

Vertical tail volume coefficient Distance between the quarter chord position of the mean aerodynamic chord of the wing and the horizontal tail

0.6 m

Distance between the quarter chord position of the mean aerodynamic chord of the wing and the vertical tail

0.6 m

Table 14: Input parameters of PA for the constraint analysis algorithm.

Important output parameters for selected wing and power loading Wing loading

10.1 kg mâ&#x2C6;&#x2019;2

Wing area

0.2024 m2

Wing span

1.0552 m

Wing mean aerodynamic chord

0.1918 m

Horizontal tail area

0.0388 m2

Horizontal tail span

0.2493 m

Horizontal tail mean aerodynamic chord

0.1558 m

Vertical tail area

0.0142 m2

Vertical tail span

0.1193 m

Vertical tail mean aerodynamic chord

0.1237 m

Wing mass

0.2020 kg

Motor mass

0.0626 kg

Battery mass

0.2700 kg

8

Empty mass Empty mass includes wing, motor and battery mass

1.1445 kg

Mission mass

2.0445 kg

Takeoff power requirement Minimum drag coefficient

188.1W 0.0293

Table 15: Output parameters for PA evaluated using constraint analysis algorithm.

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Important input parameters Maximum lift coefficient

1.5

Propulsive efficiency

0.6

Wing aspect ratio

8

Horizontal tail aspect ratio

4

Vertical tail aspect ratio

1

Vertical tail taper ratio

0.5

Horizontal tail volume coefficient

1.2 0.08

Vertical tail volume coefficient Distance between the quarter chord position of the mean aerodynamic chord of the wing and the horizontal tail

0.9 m

Distance between the quarter chord position of the mean aerodynamic chord of the wing and the vertical tail

0.9 m

Table 16: Input parameters of MSA for the constraint analysis algorithm.

Important output parameters for selected wing and power loading Wing loading

7.7 kg mâ&#x2C6;&#x2019;2

Wing area

0.2900 m2

Wing span

1.5230 m

Wing mean aerodynamic chord

0.1904 m

Horizontal tail area

0.0713 m2

Horizontal tail span

0.5425 m

Horizontal tail mean aerodynamic chord

0.1356 m

Vertical tail area

0.0380 m2

Vertical tail span

0.1981 m

Vertical tail mean aerodynamic chord

0.2054 m

Wing mass

0.2933 kg

Motor mass

0.0453 kg

Battery mass

0.2160 kg

9

Empty mass

1.5646kg

Mission mass

2.2366 kg

Takeoff power requirement Minimum drag coefficient

136.2 W 0.0293

Table 17: Output parameters for MSA evaluated using constraint analysis algorithm.

4.7

Stability and Trim

Aircraft static and dynamic stability are very important, especially for the MSA which has to fly both empty and with payload.

This results in a varying CG. Before the CAD models were developed an initial static stability analysis for both aircraft was 9 Empty

mass includes wing, motor and battery mass

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carried out. Since one of the design parameters for the constraint analysis algorithm is the distance between the quarter chord position of the mean geometric chord of the wing and the horizontal tail as well as the horizontal tail volume coefficient, another algorithm was developed which determines the stick-fixed neutral point of the wing-tail configuration for both aircraft.

The algorithm not only determines the stick-fixed neutral point but also the required position of the CG for a desired static margin and the required incidence angles of the wing and the tail such that the aircraft is trimmed at zero angle of attack with respect to its fuselage reference line and with zero elevator deflection during cruise conditions. This approach is similar to the one described by Phillips,10 with the difference being that the algorithm additionally finds the incidence angles. It was verified by the XFLR5 software, however both the algorithm and the XFLR5 software did not take into account the effect of the fuselage on the stick-fixed neutral point. Therefore an initial conservative value for the static margin of both aircraft was used as introducing the fuselages will reduce it. A more precise refinement of the CG position with respect to the neutral point will be carried out after flight testing.

The benefit of performing this analysis is that if the CG location for a particular static margin is known, for both the PA and MSA, then it can be compared with the CG from the CAD models and changes can be made if needed. An example of such change is when the CAD models of the PA and MSA were developed and the actual centers of gravity determined, it was found out that the required CG position and the actual of the MSA were in considerable disagreement. A design change was made, the wing of the MSA was moved slightly aft and the issue was resolved. After desired static stability results were achieved dynamic stability analysis were carried out with the XFLR5 software.

4.7.1

Static Stability

Table 18 shows the calculated location of the CG for desired static margin for the PA. The results from the static stability algorithm were verified against results from XFLR5. Slight differences can be noticed between the incidence angles for the PA obtained by the algorithm and XFLR5. This is due to the fact that the algorithm uses lifting line theory which is not very good for low aspect ratio wings in contrast with the vortex lattice method used by XFLR5. For the MSA the results were in a good agreement. The neutral point locations estimated by both methods agreed well.

Static stability parameters Static margin (%MAC) Calculated distance from c/4 position of the MAC of the wing to center of gravity Calculated distance from c/4 position of the MAC of the horizontal tail to center of gravity Neutral point location Wing incidence angle (Static stability algorithm) Horizontal tail incidence angle (Static stability algorithm) Wing incidence angle (XFLR5) Horizontal tail incidence angle (XFLR5)

Value 10% -0.014 m 0.586 m -0.033 m 1.7 ◦ -0.2 ◦ 2.1 ◦ -0.7 ◦

Table 18: Static stability parameters for PA. Note that all distances are measured positive aft of the center of gravity.

10 W.

Phillips. Mechanics of Flight. Wiley, 2009.

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2 Lh

α δe

1.5 1 0.5

Lift (N)

Angle (deg)

5 4.5 4 3.5 3 2.5 2 1.5 1 0.5 0 -0.5 -1 -1.5 -2 -2.5 -3 -3.5 -4 -4.5 -5 10

0 -0.5 -1 -1.5

15

20

25

30

-2 10

15

Airspeed (m s -1)

20

25

30

Airspeed (m s -1)

Figure 13: (a) required angle of attack and elevator deflection at various airspeed and (b) lift generated by the tail for PA.

0.04 0.02

Cm

0 -0.02 -0.04 -0.06 -0.08 -5

0

5

10

Angle of attack (deg)

Figure 14: Pitching moment versus angle of attack (Verification of static stability and trim algorithm results for production aircraft with XFLR5 software.)

Similarly, Table 19 shows the calculated location of the CG for desired static margin for the MSA. Static stability parameters Static margin (%MAC) Calculated distance from c/4 position of the MAC of the wing to center of gravity Calculated distance from c/4 position of the MAC of the horizontal tail to center of gravity Neutral point location

Value 10% -0.083 m 0.817 m -0.103 m

Wing incidence angle (Static stability algorithm)

-0.2 ◦

Horizontal tail incidence angle (Static stability algorithm)

-1.9 ◦

Wing incidence angle (XFLR5) Horizontal tail incidence angle (XFLR5)

0◦ -1.7 ◦

Table 19: Static stability parameters for MSA. Note that all distances are measured positive aft of the center of gravity.

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2 Lh

ι δe

1.5 1 0.5

Lift (N)

Angle (deg)

5 4.5 4 3.5 3 2.5 2 1.5 1 0.5 0 -0.5 -1 -1.5 -2 -2.5 -3 -3.5 -4 -4.5 -5 10

0 -0.5 -1 -1.5

15

20

25

30

-2 10

15

Airspeed (m s -1)

20

25

30

Airspeed (m s -1)

Figure 15: (a) required angle of attack and elevator deflection at various airspeed and (b) lift generated by the tail for MSA.

0.1

0.05

Cm

0

-0.05

-0.1

-0.15 -5

0

5

10

Angle of attack (deg)

Figure 16: Pitching moment versus angle of attack (Verification of static stability and trim algorithm results for MSA with XFLR5 software).

4.7.2

Dynamic Stability

XFLR5 software was used again to analyse the dynamic stability of both the PA and MSA. The analysis did not include the stability effects of the aircraft fuselages.

15 15

Dutch Roll mode Roll mode Spiral mode Short period mode Phugoid mode

10

10

5

5

0

0

-5

-5

Dutch Roll mode Roll mode Spiral mode Short period mode Phugoid mode

-10

-10

-15 -15 -40

-35

-30

-25

-20

-15

-10

(a) Root locus plot for PA

-5

0

-40

-35

-30

-25

-20

-15

-10

-5

0

(b) Root locus plot for MSA

Figure 17: Root locus plots for PA and MSA.

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Phugoid mode Damping ratio

Empty

Phugoid mode

0.004

Damping ratio

Ncomp = 2

0.002

0.005

Frequency

0.49 rad s

Frequency

0.80 rad s

0.62 rad s−1

Period

12.69 s

Period

7.89 s

10.06 s

Empty

Ncomp = 2

Short period mode Damping ratio

−1

Empty

Empty

−1

Short period mode

0.602

Damping ratio

0.67

0.52

Frequency

9.24 rad s

Frequency

15.38 rad s

17.12 rad s−1

Period

0.68 s

Period

0.41 s

0.37 s

Empty

Ncomp = 2

0.02 s

0.02 s

Empty

Ncomp = 2

Roll mode Time to half amplitude Dutch roll mode Damping ratio

−1

Empty

−1

Roll mode

0.02 s

Time to half amplitude

Empty

Dutch roll mode

0.168

Damping ratio

0.289

0.216

Frequency

10.00 rad s

Frequency

7.67 rad s

8.14 rad s−1

Period

0.63 s

Period

0.77 s

0.77 s

Spiral mode Time to double amplitude

−1

−1

Empty

Spiral mode

Empty

Ncomp = 2

7.85 s

Time to double amplitude

7.07 s

9.14 s

Table 20: PA mode properties.

Table 21: MSA mode properties.

From the preliminary dynamic stability analysis both aircraft were observed to be stable (level 1 performance) in all modes but the spiral. Making their spiral mode stable required introducing dihedral to the wings, however to keep the structural complexity to a minimum, which is also necessary for mass optimization, the spiral mode was left unstable. Due to the characteristics of the spiral mode, its large time to double amplitude, the pilot can easily prevent the excitement of this mode.

4.8

Propulsion

In order to design an optimal propulsion system for both aircraft, the following criteria were established: • Must be able to provide sufficient thrust for take-off.

• Weight of the whole propulsion system should be minimal, especially batteries, to give the best performance and RAC. • Propulsion system should be optimised for each mission. Especially important for the MSA, which must perform well in two missions with different requirements.

• Propellers must be within ground clearance limits. The PA propeller diameter is constrained by the cross-sectional size of the MSA fuselage, due to the chosen configuration.

• Both aircraft must be able to achieve the design speed of 20 m s, for safe flight in gusty conditions common in Wichita.

Experience gained from the previous inaugural years entry supplied the team with valuable information. In 2015 our propulsion system had struggled during the most demanding mission, consequently higher safety factors were used this year in addition to a more rigorous testing program. The Aeronaut CAMcarbon Power-Prop propellers were found to be exceptionally good as the larger diameter propellers produced comfortably more thrust than other propellers of the same size, so these would be predominantly tested where large diameter is required. RC Tiger Motors, T-Motors, were proven to be very reliable and efficient and were accordingly given extra consideration during the motor selection process.

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4.8.1

Motor and Propeller Selection

Several stages were used in putting together propulsion set-ups. Firstly, a selection of suitable motors was found, then a range of propellers and finally the ESC was selected. Power requirements and battery demands were considered throughout.

The process in full is given below.

• Thorough research was conducted to narrow down the motor choice to tried and tested, high performing brands.

T-motor, Hacker, AXI and Neu Motors were the leading contenders.

• To narrow these down further to specific models, a mixture of empirical data, theoretical calculations, power system selection software (drivecalc and motocalc mainly), manufacturer data, first-hand experience and user reviews were used.

• Further analysis on the shortlisted motors was carried out to pair them with a range of suitable propellers, ensuring the system would provide enough thrust for take-off, capable of desired speed and with reasonable power requirements.

• Electronic Speed Controllers (ESCs) were selected to support the maximum expected current draw from the motor, plus an extra 20%. A Battery Eliminator Circuit (BEC) was not required as separate batteries are used for propulsion and control, however including it did not always affect the weight, which was also a key consideration.

• Initial static thrust tests were followed by preliminary flight tests to confirm performance across the whole flight envelope and complete missions.

Based on the initial performance numbers and constraints, the type of motor and rough specifications were set. It was decided that a brushless out-runner motor would be used for both aircraft, due to their high efficiency, simplicity in using direct-drive and potential weight savings from not requiring a gearbox.

For the MSA, a KV rating (no-load RPM per volt) of between 350 and 700 was found to be most suitable, since there was little to no constraint on propeller size this would allow a large diameter, higher efficiency propeller to be used. The PA propeller would be constrained to a maximum diameter of 11 (inches), so a KV in the range 800 to 1200 would be used.

Calculations showed that propeller pitch would have to be greater than 10 inches for the MSA, and 6 inches for the PA, to achieve the desired design speeds. These, and the other requirements and constraints to be considered are shown in Table 22.

Aircraft

Maximum Power Requirement (W)

KV Range (rpm/V)

Maximum Propeller Diameter (in)

Minimum Propeller Pitch (in)

Minimum Take-off Speed (m s−1 )

Design Speed (m s−1 )

Maximum Ground Roll (m)

PA

188.1

800-1200

10

6

12.7

20

30

MSA

136.2

350-700

15

10

11.09

20

30

Table 22: Propulsion requirements and constraints.

Two motors were selected for each aircraft, which were to be tested with a combination of propellers and voltages to optimise the system and verify the data and calculations. Aeronaut CAMCarbon Power and APCE propellers were tested due to their proven high quality and having them already available. These are shown in Table 23, along with some of the propeller sizes to be tested. Initially static thrust tests were performed, and static thrust and pitch speed used as measures to ensure take-off could be achieved.

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Aircraft

Motor

KV (RPM/V)

Weight (g)

Max. Continuous Current (A)

Propellers

PA

T-Motor MT 2216

900

75

20

10x8, 10x7, 9x7, 8x6

AXI 2808/24

1190

76

22

10x7, 9x6, 9x5, 8x6

T-Motor MN 4010

370

112

20

15x12, 15x10, 14x10

T-Motor MN 3508

580

82

18

14x10, 13x10, 13x9

MSA

(in. x â&#x2014;Ś )11

Table 23: Shortlisted motors and propellers tested using Hobbywing Platinum 30A-OPTO Pro ESC.

4.8.2

ESC Selection

During testing the Hobbywing Platinum 30A-OPTO Pro ESC was used, in case of higher than expected current draw. The ESCs considered for actual flight are shown in Table 24. Lightweight designs were favoured that could handle the current draw from motor.

ESC

Current Rating (A)

Weight excluding wires (g)

BEC

Castle Creations Talon 25, 25V 25AMP ESC, with BEC

25

8

Linear

T-Motor Air 20A ESC

20

14

None

Jeti HiCopter 30A MultiCopter ESC

30

28

None

MayTech Harrier-eXTrem 25A ESC

25

25

Linear

DYS BLHeli ESC 20A Opto

20

7.5

None

Afro ESC 30Amp Multi-rotor Motor Speed Controller

30

17

Linear

Table 24: Shortlisted ESCs.

4.8.3

Battery Selection

For each aircraft and each mission the battery must be capable of supplying adequate power to the propulsion system for take-off and the rest of the mission. Collaboration with the propulsion team was required to ensure that sufficient voltage would be supplied to meet the propulsive requirements, without leading to excess battery mass, and that the current draw was not excessive. NiMH or NiCd cell chemistry could be used, but NiMH was an easy choice due to a greater energy density, which reduces mass.

NiMH cells have a continuous C rating of 10, so a 2000 mAh battery can discharge continuously at 20 A, and most are also capable of a higher brief burst current over 10 C. However, discharging at high rates leads to a very noticeable performance drop which will be tested to quantify this. It therefore may be preferable to discharge at a lower rate, especially during MF2 involving two take-offs on the same battery pack. This may increase the battery mass again though, as the power demands remain the same, so either higher capacity (and heavier) batteries or a greater number of cells may need to be used. 1600 mAh and 2000 mAh cells were the best battery capacity options for the required performance and low weight necessity. 2000 mAh cells being heavier but allowing higher current draw.

The PA will likely require 10-12 1600 mAh cells based on early propulsive estimates, propulsion testing would solidify this number and then the aim is to reduce this aggressively through optimisation.

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The MSA will have sufficient power for MF1 with 10-12 1600 mAh cells as well, again with room for optimising. Since either 10C or 8C discharge rate may be preferable for MF2, the battery and overall propulsion system mass should be considered.

The prime motor and propeller options highlighted earlier are shown in Table 25 with the battery configuration to compare battery mass. The overall propulsion system mass is estimated from cell mass and known component mass, and given a rating out of 10 for easy comparison, 10 being the lightest and therefore best. Further optimisation is required before the exact battery mass will be known. The configuration with the lowest mass may still not be optimal, however, as flying two flights will allow a lighter MSA, this was considered before the final decision to fly twice during MF2 was made.

Number of flights

Motor and propeller

Current at take-off (A)

Discharge rate at take-off (C)

Cell capacity (mA h)

Cell number

Mass rating

2

T-Motor MN4010 370Kv with 15x10

13

8

1600

12

8.3

2

T-Motor MN3508 580Kv with 14x10

16

8

2000

10

6.6

2

T-Motor MN3508 580Kv with 14x10

16

10

1600

12

8.9

1

T-Motor MN3508 580Kv with 14x10

16

10

1600

10

10

Table 25: Battery mass comparison.

Battery selection was refined from the data gathered in previous years, with more emphasis on optimizing the power system for mass for both aircraft. Previous testing had shown the Overlander SubC Premium Sport batteries to be an excellent choice, so these were again used and tested. This years testing would be focused on getting as much out of each battery as possible, by simulating missions with realistic propulsion set-ups and monitoring the state of the battery throughout each discharge cycle. Main results were focused on power available, although voltage in each cell, current draw and thrust produced were also monitored. Careful and realistic tests would allow the team to extract the maximum possible power from each battery to reduce the battery mass as much as possible.

4.9 4.9.1

Electronics Servo Selection

Servo selection was simplified for the team this year as the previous constraints on the number of servos on each aircraft was no longer present. Also, due to the scaling of the aircraft, most servos now fall in the 9 g class. Finding appropriate servos thus came down to a simple validation process to find those that had the most desirable specifications. From past experience of the team, it was found that the HobbyKing HKSCM12-5 Single Chip Digital Servo provides one of the best compromises for the sought after criteria.

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5 5.1

Detail Design Aircraft Dimensions Wing

Wing

Airfoil

NACA 4412

Airfoil

NACA 4412

Span

1.0552 m

Span

1.5230 m

Mean geometric chord

0.1918 m

Mean geometric chord

0.1904 m

Area

0.2024 m

Aspect ratio

2

5.5

0.2899 m2

Area Aspect ratio

8 1

Taper ratio

1

Taper ratio

Incidence angle

◦

Incidence angle

0

Horizontal tail

0◦

Horizontal tail

Airfoil

NACA 0012

Airfoil

NACA 0012

Span

0.2493 m

Span

0.5425 m

Mean geometric chord

0.1558 m

Mean geometric chord

0.1356 m

Area

0.0735m2

Area

0.0388m

Aspect ratio

2

1.6

Aspect ratio

4 1

Taper ratio

1

Taper ratio

Incidence angle

◦

Incidence angle

0

Vertical tail

-3.5 ◦

Vertical tail

Airfoil

NACA 0012

Airfoil

NACA 0012

Span

0.1193 m

Span

0.1981 m

Mean geometric chord

0.1237 m

Mean geometric chord

0.2054 m

Area

0.0385m2

Area Aspect ratio Taper ratio

0.0142 m2 1 0.5

Table 26: Final PA design parameters.

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Aspect ratio Taper ratio

1 0.5

Table 27: Final MSA design parameters.

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5.2

Aircraft Weights and Balance Component

Mass

CG Location

(gram)

X (mm)

Y (mm)

Z (mm)

Wings

146

-237

2

-20

Fuselage

189

-255

0

41

Propeller and spinner

21

48

0

15

Tail

47

-852

0

0

Main Landing Gear

95

-203

0

54

Tail wheel

14

-915

-4

25

Motor

100

15

0

15

Battery

329

-84

0

17

Payload

1000

-225

0

55

Receiver

18

-80

-14

73

Low Voltage Alarm

3

-94

17

77

Table 28: PA weight balance.

Component

Mass

CG Location

(gram)

X (mm)

Y (mm)

Z (mm)

Wings

155

-322

2

-25

Fuselage

239

-441

0

45

Propeller

36

238

0

13

Tail

45

-976

0

-43

Main Landing Gear

94

-207

0

168

Tail wheel

13

-1045

-4

248

Motor

100

202

0

-13

Battery

266

-26

0

83

Payload 1 PA

900

16

-101

450

Payload 2 PA Fuselage

657

5

-119

380

Payload 3 PA Wing

243

46

-54

638

Receiver

18

-72

0

85

Low voltage alarm

3

-86

-14

90

Table 29: MSA weight balance.

The reference weights coordinate systems for both the PA and MSA aircraft were placed at the beginning of the fuselage booms of the PA and MSA with the z-axis pointing aft, the x-axis pointing up and the y-axis pointing in the direction of the port wing of both aircraft. The PA and MSA components and their masses and center of gravity locations are shown in Tables 28 and 29.

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5.3

Aircraft Flight Performance Parameters

360˚ Turn

PA 14 m/s MSA 14 m/s

PA 20 m/s MSA 20 m/s Cruise 1000ft (300m) PA 14 m/s MSA 14 m/s 180˚ Turn

Takeoff

PA 16 m/s (R/C 1.26m/s) MSA 16 m/s (R/C 1.26m/s) Climb

Runway 100ft (30m)

Not to scale

Figure 18: PA and MSA flight performance parameters.

Mission 1

Mission 2

Mission 3

0.24

0.32

0.42

8.0

9.2

10.6

9.6

11.1

12.7

Best L/D speed, m s

10.0

11.9

15.9

Sustained turn load factor

2.5

2.5

1.7

CL (cruise) −1

Stall speed, m s

−1

Takeoff speed, m s

−1

Table 30: Mission performance parameters.

As the PA has greater wing loading (10.1 kg m−2 ) than the MSA (7.7 kg m−2 ), the MSA will tend to have superior sustained turn performance, therefore superior sustained turn load factor according to Phillips.12

5.4

Aircraft Mission Performance Mission 1

Mission 2

Mission 3

Takeoff time, s

4.3

8.6

3.8

Climb time, s

7.9

15.8

7.9

180 turns time, s

12.0

8.0

18.8

360 turns time, s

12.0

8.0

18.8

Cruise time, s

82.5

37.5

82.5

-

67.9

-

118.7

145.8

131.7

Taxi and landing, s Total time, s

Table 31: Mission performance estimates. Note that loading and unloading time is not included in time estimates for Mission 3.

12 Phillips,

Mechanics of Flight.

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5.5

Aerodynamic Analysis

Additional aerodynamic analysis was performed for the Production Aircraft using Computational Fluid Dynamics methods.

The main objective for the simulation was to investigate the drag increase due to the separated flow behind the fuselage.

Star-CCM+ software was used for simulations.

The PA model was placed in a bullet-shaped domain. It excluded the wing to decrease the simulation time, and symmetry condition was used so that only half the domain needed to be calculated. An alligned mesh was created using the prism layer mesher. 22 prism layers were used to capture the boundary layer correctly, the total thickness being 5 mm. 0.01 m base size was used, with finer refinement around the edges and in the wake area and coarser refinement in areas further away from the surface. The mesh resulted in 0.9M cells and can be seen in Figure 19.

Figure 19: Mesh for the PA CFD simulation.

16 m sâ&#x2C6;&#x2019;1 reference velocity was selected, and kâ&#x2C6;&#x2019; turbulence model was used for the simulation. This model is recommended over the regular Spalart-Allmaras when bigger wake areas are expected. The simulation reached convergence after 600 iterations. The high order of residuals ( 10â&#x2C6;&#x2019;4 â&#x2C6;&#x2019; 10â&#x2C6;&#x2019;3 ) suggested that for more accurate results, an improvement in meshing must be done. However, for the purpose of this simulation it was decided to be accurate enough. It has to be noted that the effects of the propeller were not included in the simulation due to a significant rise in complexity of the model.

Drag results from the simulation are presented in Table 32. The values from preliminary design stage are shown there for comparison.

CD0

Landing Gear CD0

Tail CD0

Fuselage CD0

Fuselage (aft wake only) CD0

CFD Analysis

0.0362

0.0137

0.0060

0.0165

0.0073

Preliminary Design

0.0152

0.0030

0.0048

0.0086

-

Table 32: Drag coefficient comparison from the CFD and the preliminary design stage.

As expected, the fuselage drag was highly underestimated in the preliminary design, as no correction for the wake behind the fuselage was included. The wake is estimated to almost double the fuselage drag. An unexpectedly high drag value for the landing gear was observed as well, even though in the model the landing gear was streamlined. To investigate this, wake was visualized with two turbulent kinetic energy isosurfaces at 5 J kgâ&#x2C6;&#x2019;1 (blue) and 10 J kgâ&#x2C6;&#x2019;1 (red) in Figure 20.

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Figure 20: Turbulent Kinetic Energy iso surfaces at 5 J kgâ&#x2C6;&#x2019;1 (blue) and 10 J kgâ&#x2C6;&#x2019;1 (red).

Four biggest areas of separation are observed with this wake identification method. High separation areas behind the fuselage and the landing gear wheel were expected, as high surface curvatures occur in both cases. However, the one behind the landing gear support was not considered before and will be taken into account while further optimizing the shape of the aircraft. The wake above the fuselage was not expected as well. It seems that it is the stagnation area just in front of the motor that causes this separation. It is thought to disappear due to the rotating propeller, but shifting the motor downwards (-y direction) and including a fairing above the main boom might further decrease the drag.

To conclude, aft and fore aerodynamic fairings will be considered to half the fuselage drag. Landing gear support will be taken into account and similar CFD analysis for the MSA is being prepared.

5.6 5.6.1

Structural Characteristics Load Paths

The main structural components of the aircraft are several carbon fiber tubes.

The PA fuselage contains a 10 mm outer diameter carbon fiber tube which runs through the PA foam fairing and connects to the tail assembly. A single 10 mm outer diameter tube runs along the span of both the PA and MSA wings at 25% chord. A 12 mm outer diameter carbon fiber tube runs through the foam bulkheads of the MSA fuselage. The MSA foam bulkheads are connected with carbon stringers and with foam stringers at their corners. Additionally the carbon fiber tube runs through the forward and aft MSA fuselage foam fairings. The carbon fiber tube and stringers as well as the foam stringers, bulkheads and fairings form the main MSA aircraft load paths. The load paths for the MSA and PA are shown in Figure 21.

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Wing lift Roll wrapped carbon fibre spar 10mm OD

Thrust Roll wrapped carbon fibre spar 10mm OD

Tail lift Weight

Landing gear - foam core re wrapped with carbon fibre cloth

(a) the PA

Roll wrapped carbon fibre spar 12mm OD

Pultruded carbon fibre spar 4mm mm OD Roll wrapped carbon fibre spar 10mm m OD

Wing lift

Thrust

Tail lift Weight

Landing gear - foam core re wrapped with carbon fibre cloth

(b) the MSA

Figure 21: Load paths.

5.6.2

Structural Analysis

The main structural components of the aircraft were analyzed. An assumption was made that the carbon fibre tubes along the wing and the fuselage undergo most of the structural loading. The PA undergoes a sustained load factor of 1.7 during turns, whereas the MSA undergoes a sustained load factor of 2.5 during turn. The loading applied to the 25% chord spar of both aircraft is shown in Figure 22. Assumption was made that the 25% chord spars of the wings feature distributed loading due to the lift and point loading due to the weight of the aircraft.

Based on the loads, shear and bending 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.

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y 60 y

40 60

20 40

10 20

0

x (m)

10

1.523

-10

0

x (m) 1.055

-10

distributed load (N) shear (N) moment (Nm) point load (N)

-20 -40

distributed load (N) shear (N) moment (Nm) point load (N)

-20 -40 -60

-60

(a)

(b)

Figure 22: PA (a) and MSA (b) wing carbon fibre spar shear forces and moments resulting from loads.

As with the rest of the components, the aim was to identify the lightest, yet sufficiently stiff and strong tube. Based on the last year design which utilised carbon fibre tube, and the shear force and bending moment diagrams it was selected that roll wrapped carbon fibre tubes will be used as they are stiffer than the protruded ones. Further testing of both types of carbon fiber spars was conducted so that the optimal can be found.

5.7 5.7.1

Aircraft Systems Design, Selection and Integration Production Aircraft

Fuselage The production aircraft does not feature a fuselage in its conventional meaning. A 1 mm thick, roll wrapped carbon fiber boom of internal diameter of 8 mm runs through the longitudinal axis of the aircraft. A foam fairing, needed for carrying the Gatorade bottle, is attached to the boom. Two carbon bulkheads are implemented in the design. One bulkhead is positioned at each end of the fairing. The main purpose of the bulkheads is to limit the sliding motion of the bottle inside the fairing while the production aircraft is airborne. The nose cone is positioned in front of the foam and is made out of foam. The shape of the nose cone is designed in such a way that it can accommodate the battery pack and all electronic components needed for control and propulsion. Furthermore, an aerodynamic shape that does not impose a severe drag penalty has been utilized. The motor mount that is necessary for the attachment of the electric motor represents a carbon-foam-carbon sandwich construction is bolted to the front part of the nose cone. Figure 23 illustrates the fuselage. For more details of the aforementioned components, refer to Drawing number 1, Sheet 1/6.

Figure 23: PA fuselage.

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Wing It has been decided to use a conventional configuration for the main wing. It represents a rectangle in planform shape without any sweep, taper, mechanical or aerodynamic twist. It features 14 ribs with material removed from their central section, in order to reduce the total wing mass. The ribs have been made out of craft foam. The main spar has a circular cross-section with external diameter of 10 mm, runs from one wingtip to the other and is secured at 25 % of the wing chord. The main wing features one pair of outboard ailerons which are 30 % of the wing chord. Foam sheeting is used in order to cover the leading and trailing edges between the ribs. Figure 24 shows the wing configuration.

The central foam section of the main wing is attached to the body of the PA by using 4 nylon bolts. The design team also decided to place the main landing gear between the wing and the aircraft body. In order to reinforce this attachment 2 nylon bolts have been used. Figure 24 illustrates how the 2 nylon bolts hold the landing gear to the wing, while the other 4 are needed for the attachment between the body and the main wing.

The aileron servos are attached to the wing spar using simple brackets. In Figure 24 it can be observed that the leading edge foam sheeting is extended in order to house the servo brackets on each side of the wing.

Figure 24: PA Wing.

Empennage The tail attachment is designed to utilise 3D printing which would help overcome the challenge of fitting the empennage to the single main boom. This provided freedom as almost any shape could be achieved. The final 3D printed tail attachment is a complex part that houses both the horizontal and vertical stabilizers and the rear landing gear. It is made out of Polylactic acid, also known as PLA. Usually ABS is considered for engineering purposes but after comparing exactly the same parts using different printers and both materials, PLA proved to be stronger for this purpose. This was a surprise and probably falls down to the quality of the 3D printer itself, but that played a key decision in the choice of material.

Both stabilizers are locked in and glued to small booms coming out of the part. The part itself fits onto the aircraft main boom. The elevator and rudder are attached to the main body of the respective stabilizers via small hinges. Both of them are operated by servos located on 3D printed beds which are attached to the main boom. The connection between the servos and the elevator/rudder is done by linkages attached to control horns on each side. The design of the PA features a tail-dragger configuration, which means that the rear part of the landing gear is located at the aft end of the aircraft. In this case it is a single wheel connected via a bended steel rod to a hole through the 3D printed attachment. Two shaft collars are placed on the top and bottom where the bending steel rod pierces the PLA part of the empennage. This restricts unwanted movement of the rod. The empennage configuration is shown in Figure 25.

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Vertical stabilizer Rudder Horizontal stabilizer Elevator

Landing gear

Servo attachment

Tail attachment (PLA)

Fuselage spar

Figure 25: PA Tail.

Propulsion Results from static thrust tests and flight tests validated the selections considered during the preliminary stage. The T Motor MT 2216 900Kv motor paired with a 10x7 APCE propeller is chosen for the PA. It would produce sufficient thrust for take-off, while drawing current lower than other options and would achieve the best thrust/power ratio.

The Castle Creations Talon 25, 25V 25AMP ESC would be used due to its light weight and it can cope with the current draw with a good factor of safety. 10 1600mAh cells will be used as the base for battery optimization in further testing, although it is likely the battery mass can be reduced further.

5.7.2

Manufacturing Support Aircraft

Fuselage The fuselage of the MSA consists of 3 components: the nose cone, the central section and the tail cone.

The central section contains 6 evenly spaced frames that are made out of foam. Each of the frames has a square cross-sectional shape and is 10 mm thick so that the entire structure is lighter and the sub-assemblies of the PA can be transported inside. A roll wrapped carbon fiber boom of external diameter of 12 mm goes through the top section of the frames so that they are aligned properly with respect to each other. In addition to that, 8 balsa wood stringers run parallel to the main boom in order to reinforce the fuselage structure and make it easier to cover it with film at the the final stage of the manufacturing process.

The nose cone is also made out of foam and is adhesively bonded to the first frame of the central section. Square and trapezoidal foam pieces are cut out from every side of the nose cone in order to reduce its weight and also to allow access to the electronic equipment needed for propulsion and control. The motor mount plate represent a carbon-foam-carbon sandwich construction and is bolted to the front of the nose cone.

The tail cone is attached to the last frame of the central section of the fuselage and is detachable in order to provide a way to load the sub-assembles of the PA into the fuselage. Similarly to the the nose cone, it also features lightening sections that are cut out. The tail wheel is fixed to the tail cone and a separate servo rotates the wheel so that steering is possible when the aircraft taxis on the runway. Figure 26 depicts the fuselage structure. For more details, refer to Drawing number 4, Sheet 4/6.

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Figure 26: MSA Fuselage.

Wing Regarding the planform shape and internal structural layout, there is no difference between the main wings of the PA and MSA. Both feature the same design concept and similar attachment technique to the aircraft body - via 4 Nylon bolts.

The only difference is that the main landing gear is not bolted to the main wing, but to the fuselage via a carbon plate which is positioned in front of the main wing. Figure 27 shows the main wing with the Nylon bolts necessary for its attachment.

Figure 27: MSA Wing.

Empennage

The empennage features a conventional configuration. A roll wrapped carbon tube, 5 mm in external diameter,

runs along the span of the horizontal stabilizer in order to provide additional bending stiffness. In total, 2 Nylon bolts are used to bolt the horizontal stabilizer to the place where the fuselage central section and the tail cone join together. The elevator has the same planform shape and span as the elevator of the PA. The vertical stabilizer is adhesively bonded to the horizontal stabilizer and a trapezoidal shape is cut out of it, in order to reduce its weight. Furthermore, in order to ensure good bending stiffness, a 10 mm carbon fiber roll wrapped tube serves as a spar for the vertical fin. Figure 28 illustrates the empennage structure.

Figure 28: MSA Tail.

Propulsion The MSA will use the T-Motor MN 3508 580Kv with the same ESC as the PA, the Castle Creations Talon 25, 25V 25AMP ESC. Static thrust and battery discharge tests proved this to be the best motor for best performance in Missions U NIVERSITY OF G LASGOW

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1 and 2. The ESC is again ideal for expected current draw, and lightweight design is a large part of the choice. For Mission 1 a 13x10 Aeronaut CAMCarbon propeller will be used, drawing low amount of current for this less demanding mission. 10 1600mAh cells would be sufficient, with room for optimizing. For Mission 2, the optimal propulsion set up and number of flights for the MSA is 2 flights with a 14x10 Aeronaut CAMCarbon propeller, discharging at 10C for first take-off and using extra 2 cells to give 12 in total, so take-off can be achieved in the second flight. This choice is made as the team is in agreement that more weight can be saved by being able to carry the two PA subassemblies separately, than could be saved in battery mass from not requiring two flights.

5.8

Rated Aircraft Cost

Based on the weight estimations from both the PA and MSA CAD models, shown in Tables 28 and 29, the rated cost for Ncomp = 2 was calculated to be 0.872, therefore consistent with the score and sensitivity discussion in the conceptual design.

However after manufacturing and flight testing the aircraft rated cost can change due to manufacturing tolerances or required modifications.

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1064 41.90

SCALE 1:10

293 11.53

195 7.68 247 9.72

81 3.19

249 9.81 109 4.29 120 4.72 192 7.55 55 2.17

NOTE:Distance from rib to rib along the whole wing is 55mm (2.17inches)

1023 40.26

47 1.84 10 0.39

119 4.70

157 6.19 731 28.79

323 12.72 441 17.36

157 6.19

Cessna/Raytheon Missile Student Design/Build/Fly competition

TITLE: 3-View Drawing

(Production Aircraft)

SIZE

A

UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 1 MM AND THEN IN INCHES

SCALE:1:12 (Unless specified)

Sheet 1 of 8

ITEM NO. 1

PART NAME

MATERIAL

QTY.

Carbon Fiber Carbon Fiber

1

2

4

Fuselage Main Boom Bulkhead Gatorade Bottle Fairing Fairing Lid

Crafting Foam

1

5

Motor Mount

6

Electric Motor Propeller Nose Cone - Bottom Fairing Wing Attachment Section Wing Tip Rib Leading Edge Wing Skin Wing Rib Wing Leading and Trailing Edges Wing Landing Gear Support

Carbon and Foam Aluminium Carbon FIber

2 3

7 8 9 10

11 12 13

14

17

Servo Main Landing Gear Main Landing Gear Wheel - 40 mm

18

Main Landing Gear Wheel Rod

19

25

Horizontal Stabilizer Vertical Stabilizer Rudder Elevator Empennage Attachment Part Rear Landing Gear Rear Landing Gear Wheel - 30mm

26

Collar Restraint - 3mm

15 16

20 21 22 23 24

Crafting Foam

Crafting Foam

2 2 14 2 2

Crafting Foam

Plastic Carbon Fiber Plastic 2014Aluminium Alloy

4

Crafting Foam

1

Crafting Foam

1

Crafting Foam

1

Crafting Foam

2

PLA Bending Steel Plastic 2014Aluminium Alloy

1

Plastic Plastic Plastic NiMH PLA

1

4

Crafting Foam

1 1

Propeller Spinner

36

Servo Battery

NiMH

1

37

Main Wing Spar

Carbon Fiber

1

31 32 33

30 22

13 15 16 32

19

1 10

5

7

17

6

27

4 3

3

35

30

11

1

1

36

35

8

28

29

18

Cessna/Raytheon Missile Student Design/Build/Fly competition

33

TITLE:

1

6 1

25

14

1

34

29

31

2

1

24

37

2

Carbon Fiber Plastic 2014Aluminium Alloy

28

31

1

Gatorade Bottle Receiver Low Voltage Alarm Motor Battery Pack Servo Attachment Brackets Nose Cone - Top Fairing Plate for electronic Components Nylon Bolt

27

26 23

1 1

13

15

1

Crafting Foam

Crafting Foam

20

9 12

1

1

Crafting Foam

34

1

Crafting Foam

Crafting Foam

21

SIZE

B

Structural Arrangement (Production Aircraft) UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 2 MM AND THEN IN INCHES

SCALE:1:5 (Unless specified)

Sheet 2 of 8

Aircraft Boom Motor Battery Pack Receiver Motor

Prop Spinner

Motor Compartment Scale 1:2 Low Voltage Alarm

Aileron Servo Scale 1:2

Servo Boom Servo Attachment

Servo Battery Pack

Elevator and Rudder Servos Scale 1:1

Cessna/Raytheon Missile Student Design/Build/Fly competition

TITLE:

Side View Scale 1:7 .

Front Landing Gear

Rear Landing Gear

SIZE

B

Systems Layout (Production Aircraft) UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 3 MM AND THEN IN INCHES

SCALE:1:10 (Unless specified)

Sheet 3 of 8

Foam Hatch

Gatorade Bottle

Gatorade Storage Scale 1:2

NOTE: Payload Casing and Front Landing Gear have been ommited for clarity

Payload Casing

Bulkhead

Gatorade Bottle Side View SCALE 1:2

Payload Casing Scale 1:2

Cessna/Raytheon Missile Student Design/Build/Fly competition

NOTE: Foam hatch is detachable and is securely TITLE: Payload Layout fixed during flight. The Payload casing acts as a (Production Aircraft) fuselage for the production aicraft SIZE

B

UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 4 MM AND THEN IN INCHES

SCALE:1:5 (Unless specified)

Sheet 4 of 8

1492 58.74

526 20.71

357 14.07 472 18.58

543 21.36

120 4.72 136 5.34 220 8.66 190 7.50 55 2.17

NOTE:Distance from rib to rib along the whole wing is 55mm (2.17inch

12 0.47 5 0.20

33 1.30

1342 52.82 165 6.50

132 5.20 185 7.30

220 8.66 50 1.97

136 5.34

801 31.52

Cessna/Raytheon Missile Student Design/Build/Fly competition

TITLE: 3-View Drawing

(Support Aircraft)

SIZE

A

UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN MM AND THEN IN INCHES 5

SCALE:1:15 (Unless specified)

Sheet 5 of 8

ITEM NO.

PART NAME

MATERIAL

QTY.

1 2

Carbon Fiber Crafting Foam PLA Balsa

1 1

4

Main Boom Skeletal Structure Motor Bulkhead Stinger

5

Wing Attachment Plate

Crafting Foam

1

6

Crafting Foam Aluminium Carbon Fiber

1

8

Vertical Stabiliser Motor Propeller

9

Main Landing Gear Wheel - 80mm

Plastic

2

10

Rear Landing Gear Wheel - 40mm

Plastic

1

11

Wing

Crafting Foam

1

12

Main Landing Gear

Carbon Fiber

1

13

Motor Battery

NiMh

1

14

Low Voltage Alarm

Plastic

1

15

Plastic Plastic Crafting Foam

1

17

Receiver Elevator And Rudder Servos Fuselage Corners

18

Fuselage Front Fairing

Crafting Foam

1

19

Plastic Crafting Foam

2

20

Aileron Servo Rear Servo Accomodation

21

Nylon Bolt

Plastic

7

22

25

Collar Restraint - 3mm

Crafting Foam Crafting Foam Crafting Foam 2014Aluminium Alloy

1

24

Horizontal Stabiliser Elevator Rudder

26

Tail Landing Gear Mount

Bending Steel

1

27

Foam Hatch

Crafting Foam

1

28

Propeller Spinner

Aluminium

1

29

Electronics Plate

Carbon Fiber

1

30

Servo Batteries

1

31

Wheel Rod

NiMh 2014Aluminium Alloy

32

Wing Tip

Crafting Foam

2

33 34 35

Wing Rib Aileron Wing Boom

Crafting Foam Crafting Foam Carbon Fiber

22 2 1

36

Servo Boom Attachment

PLA

2

3

7

16

23

6 19

34

36

21

16

1

16

21

27

24

23

8

1

35

22

1

2 4

20 11 25

26

2

10

12 19

2 14

36

33 1

32

1

2 4

2 1

31

3

2

23

9

25

17

5

Cessna/Raytheon Missile Student Design/Build/Fly competition 28

8

7

3

18

13

29

30

15

TITLE: Structural Arrangement

(Support Aircraft)

SIZE

B

UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 6 MM AND THEN IN INCHES

SCALE:1:8 (Unless specified)

Sheet 6 of 8

Low Voltage Alarm Servo Battery Motor Battery Main Boom Motor Propeller

Motor Compartment Scale 1:3 Prop Spinner Aileron Servo Scale 1:2

Servo Boom Servo Attachment

Receiver

Elevator and Rudder Servos Scale 1:2

Cessna/Raytheon Missile Student Design/Build/Fly competition

TITLE:

Side View

SIZE

B

Front Landing Gear

Rear Landing Gear

Systems Layout (Support Aircraft) UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN MM AND THEN IN INCHES 7

SCALE:1:10 (Unless specified)

Sheet 7 of 8

Fuselage

Aft View Scale 1:10

Main Boom

Aircraft Boom

Stinger

NOTE: Foam hatch is detachable and is securely fixed during flight.

Stinger

Foam Structure Payload Casing

NOTE: The Payload casing acts as a fuselage for the support aircraft. It is designed in such a way as to fit a large portion of the support aircraft in order to limit the number of flights.

Foam Hatch

Cessna/Raytheon Missile Student Design/Build/Fly competition

TITLE: SIZE

B

Payload Layout (Support Aircraft) UNLESS OTHERWISE SPECIFIED DWG. NO. DIMENSIONS ARE GIVEN IN 8 MM AND THEN IN INCHES

SCALE:1:5 (Unless specified)

Sheet 8 of 8

6

Manufacturing Plan

6.1

Process

The accurate and lightweight construction of a model aircraft that closely follows the design intent is crucial for a competitive aircraft. There is a wide variety of manufacturing methods available for the creation of a model aircraft which result in a lightweight and stiff structure. However some rely on a great deal of experience while others may be too costly. The typical methods of model aircraft construction are outlined below: 1. Traditional Balsa Build This entails using a mixture of very light plywood and various thicknesses of balsa to achieve a former and longeron structure, with some areas being skinned using very thin balsa (around 1/16 in). The resulting structure is then covered with a film. The most convenient method of carrying out this manufacturing is to cut as many shapes as possible using a CNC laser cutter while keeping hand shaping to a minimum. It is possible to create semi complex shapes with a mixture of CNC cutting and hand finishing.

2. Foam Cored Structure This is a feature typically seen on homebuilt and home-designed light aircraft due to the ease of build and very accurate final dimensioning. The foam core structure allows a very accurate wing manufacture where it is easy to incorporate twist and taper through the use of a CNC hot wire cutter. Depending on the stresses a particular area of the design will endure, composite materials such as glass or carbon fibre are applied to the outer surfaces creating a very rigid structure in both bending and twisting through multi-axial fibres. Complex shapes are possible with hand finishing.

3. Carbon Fibre Shell To manufacture an aircraft from a composite shell, it is required to create durable female moulds of all parts so that vacuum infusion may be utilised. A CNC router is required to create the moulds. The resulting parts have an excellent fibre resin ratios and as such are lightweight and may be tailored to the calculated stresses. The surface finish is also excellent if performed correctly with no skinning required.

The decision for the method of manufacturing for this year has been specifically considered for each component. Through hand calculation and CAD modelling it was found that a wing with a carbon spar and foam structure will offer a lighter solution than a foam cored wing, while it was found that a carbon fibre wrapped foam structure for the undercarriage would be the most appropriate.

6.2 6.2.1

Aircraft Subsystems Wing

Both the PA and MSA wings feature a foam structure. The parts from the CAD models of both aircraft were exported to a CNC hotwire machine, with which they were manufactured. The carbon fiber spar running along the wing provides resistance to bending and torsion, and several ribs were reinforced with thin glass fiber cloth. The same procedure was used for the trailing edge.

6.2.2

Fuselage

The fuselage of the PA was made entirely out of foam fairings, which were also cut by the CNC. They are attached to the carbon fibre spar running along the length of the fuselage and are reinforced at several places with thin glass fiber cloth.

The MSA fuselage features foam bulkheads attached to the carbon fibre spar running across the fuselage. The bulkheads are connected with carbon fibre stringers. The aft and forward firings of the MSA fuselage were cut in a similar manner to the ones of the MSA. U NIVERSITY OF G LASGOW

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6.2.3

Tail

The tail for both the PA and MSA were cut out of solid foam; this was chosen due to the rapid prototyping speed. However, it was decided that later versions of the tail will be made out of ribs.

6.2.4

Landing Gear

Landing gear for both the PA and MSA was created by wet layup of carbon fibre cloth on a pre cut landing gear foam shape.

The prototype wing and tail for the PA and fuselage for the MSA are shown in Figure 29 below.

(a)

(b)

(c)

Figure 29: PA and MSA prototype manufactured parts.

6.3

Schedule

The manufacturing schedule for the PA and MSA is shown in Figure 30 below.

Figure 30: Manufacturing schedule.

7

Testing Plan

The use of theoretical and computational models has its limitations as many assumptions are made to reduce the complexity, creating a need for testing. The aim is to evaluate the accuracy of the models, validate their output and verify the predicted parameters. Test results provide an indication as to whether the components are over designed or if they need further improvement, making these results invaluable for the optimization process.

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1. Propulsion testing (a) Static thrust tests will be carried out to determine optimal propeller-motor combination 2. Battery testing (a) Battery performance and available power for second takeoff will be evaluated 3. Structural testing (a) Deflection of PA and MSA wing and fuselage carbon fiber spars under load will be evaluated 4. Flight testing (a) Both the PA and MSA will be tested according to a specific flight sequence: taxiing, empty flight, payload fixture, flight with payload and landing 5. Payload fixture testing (a) Ensure that the PAâ&#x20AC;&#x2122;s sub-assemblies and the Gatorade payload will fit entirely and securely within the MSAâ&#x20AC;&#x2122;s and PAâ&#x20AC;&#x2122;s fuselages

7.1

Propulsion Testing

Propulsion testing will determine the optimal propeller-motor combination. Static thrust testing will be conducted for several propeller-motor combinations, for both the MSA and PA. Large diameter propellers with a range of pitches will be tested with higher power motors for the MSA; smaller propellers, with a focus on optimising for weight by utilising lighter motors, will be tested for the PA. The tests will be conducted using a rig assembled for this purpose, with a set of scales which will act as a load cell. Dynamic thrust testing will not be carried out due to time and budget constraints, and because of a lack of access to wind tunnel facilities. The motor will be attached to the top of the vertical arm, with the propeller blade tightly secured to the motor. The motor will be connected to an electronic speed control (ESC), which will be connected to a power analyzer which will output the current drawn, voltage across, and power of the motor. The rotational speed of the motor will be controlled by a servo tester connected to the ESC. A horizontal arm will transmit thrust force to the electronic scales although there will be adjustment made for the slightly differing length of the rig arms, caused by different mounting points being used for the different motors. Additionally, the rotational speed of the propeller will be found using a laser-tachometer.

Initial testing will be conducted to find the maximum thrust values of each propeller-motor combination. Subsequent testing will be carried out at 25%, 50% and 75% power to obtain thrust for a range of power values, and thus establish thrust-power profiles.

7.2

Battery Testing

Battery tests will be performed to confirm the difference in battery mass from performing two flights, carrying components separately, compared to one flight carrying both components. To accurately compute the difference, two tests with realistic power demands will be carried out. Firstly a simulated mission with a take off at 10C discharge rate, and following from these results, the same test will be performed at 8C. The power demands will be obtained by the mission profile for which required power at different stages of the mission was determined. The power values were converted to percentages relative to the take off power. A motor, propeller and ESC combination will be selected which would draw the necessary current.

The propulsion set up will be mounted on the static thrust rig, and the power measured by a watt meter and controlled by a servo tester. The percentage of power from first flight which could be achieved on the second flight will be analyzed.

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7.3 7.3.1

Structural Testing Wing and Fuselage Spar Testing

Wing spar testing for the PA and the MSA will simulate aerodynamic loads during a 5G turn. A safety factor of 1.2 will be added resulting in aerodynamic loads to be tested for a 6G turn. The loads simulated are MTOW for the PA and the MSA. The aerodynamic loads are represented as a point load acting at the quarter chord location on the mean geometric chord of the PA and the MSA wings. The deflection of the PA and the MSA wing carbon fibre spar will be analysed. A similar approach will be used with the carbon fibre spars of the PA and the MSA fuselages. The spar will be suspended from two points, wing and horizontal tail aerodynamic centers and the aerodynamic load applied at the aerodynamic centers and the aircraft center of gravity. Deflections will be analyzed.

7.3.2

Landing Gear Testing

The landing gear testing for the PA and the MSA will simulate a load resulting during impact. The load will be 3 times the MTOW weight of the PA and MSA. The load will be applied at the point where the landing gear is attached to the fuselage, for the MSA, and to the wing, for the PA. The landing gear movement will be restrained so that only lateral deflection is observed and measured.

7.4

Flight Test

Flight testing for both the PA and MSA will be carried out in a specific sequence and a checklist for both will be developed.

The sequences goal is to validate not only the flight envelope of the aircraft, but also the taxiing and handling qualities.

1. Taxi testing (a) The MSA and PA will be taxied with different payload configurations and the ground handling qualities will be analysed 2. Empty flight tests (a) The MSA and PA will be flown empty at favourable weather conditions. The empty flight will be a low altitude one and from it the stability characteristics of both aircraft will be analysed 3. Payload fixture testing (a) It will be ensured that the PA’s sub-assemblies fit entirely and securely within the MSA’s fuselage as well as that the Gatorade bottle fits entirely and securely in the PA’s fuselage 4. Payload flight tests (a) The MSA and PA will be flown with their respective payloads. Payload mass will be changed so that at first it is lower than the one that the respective aircraft has to carry. The MSA will be flown with different payload arrangements. From this flight the stability characteristics of both aircraft will be analysed.

7.5

Testing Schedule

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7.5.1

Checklist Inspection Item Aircraft Wing Empennage Payload Batteries

Landing Gear Motor Propeller Center of Gravity

Flight Log Batteries Radio controls Control surfaces and front landing gear Motor

Fuse and Batteries Radio controls Aircraft Flight Log

8 8.1

Task INITIAL CHECKS Visually verify that the aircraft in general is in normal condition without noticeable cracks, missing parts or unsecured components No fractures present and wing is firmly in place. Control surfaces move easily.

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

Payload is properly secured and the payload is correctly located.

Make sure battery contacts and leads are in good condition for propulsion, receiver and transmitter batteries.

No visible cracks and wrapping is intact.

Batteries are fully charged.

Battery positions are appropriate for the payload carried.

Firmly in place, no fractures. Tyres spin easily and front gear turns in normal limits.

Mount is free of fractures. Securely in place.

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

Ensure CoG matches with the loaded payload. Max aft position allowed for PA is 0.223m and for MSA 0.498m from the front tip of the main boom.

PRE FLIGHT CHECKS Fill in Flight Log details Connect batteries and then connect fuse. Make sure connections are firm.

Transmitter on first, receiver on second.

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

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

POST FLIGHT CHECKS

Unplug the fuse and batteries

Receiver off first, transmitter second Visually verify that the aircraft is in normal condition without noticeable cracks, missing parts or unsecured components.

Ensure flight is properly documented. Report any discrepancies.

Performance Results Propulsion Testing Results

Two motors for each aircraft were tested for maximum thrust at a range of voltages with a variety of propellers. Voltage and amperage were considered during the tests to determine which batteries would be most appropriate and how long the batteries would be able to sustain full power. Thrust per unit Power was used as a factor to determine the overall effectiveness of the given power system, while pitch speed and static thrust were used to ensure the aircraft would be able to take-off and fly at the design speed. The best options found for each aircraft are shown in Tables 33, 34, 35, and 36.

Further testing on the batteries was carried out as the MSA motor and propeller choice would be strongly affected by second takeoff. Static thrust against power was also plotted to use as a comparative tool in final selection. They were generated by measuring the thrust at a range of powers. The graphs are shown in Figure 31, and the rig used is shown in Figure 32.

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Propeller Voltage (V)

Current (A)

Power (W)

RPM

Static thrust (N)

Pitch speed (m s−1 )

Static thrust/Power (N W−1 )

10x7

12.3

14.4

177.4

7300

7.85

21.63

0.048

11x9

11.9

18.8

224.3

5250

8.32

20.00

0.040

Table 33: T Motor MT 2216 900Kv test results.

Propeller Voltage (V)

Current (A)

Power (W)

RPM

Static thrust (N)

Pitch speed (m s−1 )

Static thrust/Power (N W−1 )

7x5

11.9

12.4

147.4

12050

4.16

25.51

0.031

8x6

12.3

20.5

252.2

10750

7.39

27.31

0.028

Table 34: Axi 2808/24 1190Kv test results.

Propeller Voltage (V)

Current (A)

Power (W)

RPM

Static thrust (N)

Pitch speed (m s−1 )

Static thrust/Power (N W−1 )

14x9

15.2

11.7

177.8

4505

11.59

17.16

0.076

15x10

15.2

13.35

202.9

4364

13.39

18.47

0.077

Table 35: T-Motor MN 4010 370Kv.

Propeller Voltage (V)

Current (A)

Power (W)

RPM

Static thrust (N)

Pitch speed (m s−1 )

Static thrust/Power (N W−1 )

13x10

12.5

13.4

167.9

4873

10.54

20.63

0.061

14x10

12.4

15.6

198.0

4563

12.46

19.32

0.071

Table 36: T-Motor MN 3508 580Kv.

9

14 T Motor MT 2216 10x7 T Motor MT 2216 11x9 Axi 2808/24 7x5 Axi 2808/24 8x6

7

T Motor MN 4010 14x9 T Motor MN 4010 15x10 T Motor MN 3508 13x10 T Motor MN 3508 14x10

12

Static Thrust (N)

Static Thrust (N)

8

6 5 4

10 8 6

3 4

2 1

2 0

50

100

150

Power (W) (a) PA Static thrust versus Power

200

250

0

50

100

150

200

250

Power (W) (b) MSA Static thrust versus Power

Figure 31: PA (a) and MSA (b) static thrust versus power for best motor-propeller combinations.

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Figure 32: Static thrust rig used for tests.

8.2

Battery Testing Results

The battery discharge test results simulating Mission 2 with two flights are shown in Figure 33, showing power demand across the mission profile. The power is shown as a percentage of the maximum achieved during flight one with 10C discharge rate. Discharging at 8C shows no drop in available power for second take-off, while discharging at 10C leads to 25% less power. This means greater initial power would be required for discharge at 10C. However discharging at 8C may increase the cell number required, so the battery mass was still compared for each case.

100

100 10C Discharge Rate 8C Discharge Rate

80 70 60 50 Take-off

Cruise

10C Discharge Rate 8C Discharge Rate

90

Percentage of Max Power (%)

Percentage of Max Power (%)

90

Taxi

40 30 20 10

80 70 60 50 Take-off

Taxi

40

Cruise

30 20 10

0

0

0

10

20

30

40

50

60

70

80

time (s)

(a) Battery discharge during first flight of mission 2

0

10

20

30

40

50

60

70

80

time (s)

(b) Battery discharge during second flight of mission 2

Figure 33: MSA Battery discharge results during Mission 2.

8.3 8.3.1

Structural Testing Results Spars

The carbon spars in the wing of the PA and MSA aircraft were tested under different load factors. The maximum deflection at the tip of the spars was measured, as demonstrated in Figure 34. The aerodynamic load was positioned at the quarter chord position of the mean geometric chord of the wing. Since the wing spar is at the quarter chord, the location of the mean geometric chord along the spar was determined. Aerodynamic load was changed appropriately to represent different load factors. Two different outer diameter spars carbon manufactured by two techniques, pultruded and roll wrapped, were

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tested. The results from the maximum deflection of each of the carbon fibre spars is shown in Table 37. From the results the roll wrapped carbon fiber tubes were selected as the deflection of the 10 OD roll wrapped was close to the 12 OD protruded, meaning that the 10 OD roll wrapped was strong enough but lighter than the 12 OD one.

10 OD Pultruded Tip Deflection (mm)

12 OD Pultruded Tip Deflection (mm)

10 OD Roll wrapped Tip Deflection (mm)

12 OD Roll wrapped Tip Deflection (mm)

2G

24

24

25

21

3G

51

46

49

43

4G

75

68

71

62

5G

105

96

98

87

6G

142

130

136

124

Load

Table 37: Carbon fiber wing spar test results.

Figure 34: Carbon fiber spar tests.

8.3.2

Landing Gear Testing Results

The loaded (3 times the MTOW) landing gear of the PA is presented in Figure 35a and the test setup for the PA landing gear drop test is shown in Figure 35b.

No visible deformation was observed during the static landing gear loading test. For the dynamic test, the PA was dropped from a height of 0.4m. The drop caused the PA landing gear to crack at its starboard bend, indicating poor impact absorption.

For the next design, increase in foam thickness at the joint is suggested to improve the impact absorption.

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(a) Static Test

(b) Drop Test

Figure 35: Landing gear tests.

8.4

Flight Testing

Three test flights for the PA were done, the main objectives being: 1. Proof of handling qualities for a very high wing loading aircraft.

2. Check for takeoff distance requirement.

3. Basic flight data gathering with the telemetry pack.

The flights were performed in Cumbernauld Airfield, which has a 62 m long tarmac runway, 8 m in width, with a grass landing possible on both sides of the runway. The following order of the test flights was kept: 1. Empty test flight.

2. Half payload test flight together with the telemetry pack.

3. Full payload flight with trajectory similar to the Mission 3 profile.

All the flights were recorded on camera. In addition, a basic telemetry system was put together to provide flight information on the PA. This was comprised of an accelerometer, a gyroscope, a barometer and a GPS receiver. Unfortunately, the barometer, accelerometer and gyro assembly were found to malfunction shortly prior to takeoff and had to be disabled.

However, the GPS antenna was operable, and real time radio telemetry could be read while the aircraft was flying. An on-board copy of the data was also present but served as a backup in the event of a failure in the real time telemetry system.

8.4.1

Results

The takeoff distance requirement was satisfied in all three flights. The duration of the three flights were 1min41s, 1min25s and 2min49s. The third flight proved the aircraft to be capable of successfully completing the Mission 3. The handling qualities, according to the pilot, were good and the aircraft was landed on the tarmac with no difficulties. No damage to the aircraft was done.

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The available data from the second flight was analyzed. In total, three high quality legs (data intervals less than 5 s) were available. The flight paths are plotted in Figure 36 with the important parameters summarized in Table 38. It has to be noted that no accurate weather information was recorded on the day of flight, hence the velocities do not include correction for the wind speed. However, mean winds were recorded by the meteorological offices in the area to be of S-SSW heading with 0-1 m s−1 velocities.

Figure 36: Flight path for the first test flight. The different colours mark the different flight legs, the white arrow marks the first received data location and markers M and V mark locations where the maximum and maximum level flight velocities were achieved.

Flight Leg

Distance (m)

Vmax (m s−1 )

Vmin (m s−1 )

Vavg (m s−1 )

Level Flight Vmax (m s−1 )

Flight Altitude (m)

γmax (◦ )

R/Cmax (m s−1 )

1

344

20.9

7.0

15.1

20.4

199

32

10.4

2

452

23.1

10.2

17.6

19.6

195

29

5.6

2

198

17.6

10.8

14.0

17.6

162

18

9.2

Table 38: Flight test data for the three flight legs. Here Vmax , Vmin and Vavg are the maximum, minimum and average flight velocities, γmax is the maximum flight path angle and R/Cmax is the maximum rate of climb.

in Figure 36, the white arrow marks the location where the data was received for the first time. Black lines mark the missing connections in between the three legs. Markers M and V mark the locations where maximum velocities and maximum velocities in level flight (flight path < 4 ◦ ) were reached, correspondingly. Several important observations can be made from the data.

Firstly, several turns were performed at turn velocities around 20 m s−1 , proving the wing strength of the PA to be sufficient not only for the design cruise speed (20 m s−1 ), but for turns as well. Also, the highest velocity reached (23.1 m s−1 ) assures that the design speed is below the flutter speed. Achievement of high flight path angles with no loss of control indicates that no non-linear effects in longitudinal stability appear. Last year, not measuring the latter two parameters induced aileron flutter and stability problems. Therefore, it was important for the team to make sure the problem does not appear this year again.

Minimum recorded flight speed was 7 m s−1 , which with mean wind correction would be around 8 m s−1 . In comparison, the design stall speed for half the flight payload was set at 9.3 m s−1 . Hence, either the design stall speed is too conservative, or the mean wind data leaves too big the error to be accurate. Therefore this comparison has to be interpreted with care.

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High climb rates indicate overpowering of the aircraft. Even with the full payload, the aircraft was seen to perform very well in climbs. Reduction in the motor power and the battery size is possible, potentially leading to further mass optimisation.

Flight test for the MSA has not yet been performed. However, important conclusions were made from the PA test flight and will be taken into account in following flights. The data rate of the telemetry system is also planned to be increased.

Figure 37: PA test flight.

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