Final report Team 5

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1. EXECUTIVE SUMMARY

1.1 Team organization

2. MILESTONE CHART

3. CONCEPTUAL DESIGN

3.1 Mission requirements

3.2 Translation into design requirements

3.3 Considered configuration

4. PRELIMINARY DESIGN

4.1 Critical design parameters

4.2 Mission analysis

4.3 Design and sizing trades

4.4 Stability and control analysis

4.5 Structural analysis

5. DETAILED DESIGN

5.1 Dimensional parameters

5.2 Weight and balance

5.3 Flight performance parameters

5.4 Mission performance

5.5 Drawing package

6. MANUFACTURING PLAN AND PROCESSES

6.1 Investigation and selection of major components and

assemblies

7. REFERENCES

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I. Executive summary

This report describes the design process used by Air-Navigation students, 5th

team to

develop an aircraft capable of winning the 2012 Design/Build/Fly Competition. The goal of the

design was to maximize the total competition score, being a combination of the report score,

Solid Works design, XFLR fuselage computations and MathCad problems.

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II. Milestone Chart

The chart below shows the progress over the time allowed, from the beginning of

November up to the presentation date. Although, sometimes we did not meet certain deadlines,

we managed to accomplish a balance between deadlines and quality.

item descrition Start date End date 2011 2011 2011 2012 2012 2012 2012 2012

Oct Nov Dec Jan Feb Mar Apr May

1 Complete

project

24.10.2011 18.05.2012

2 Conceptual

design

24.10.2011 31.10.2011

3 Preliminary

design

07.11.2011 14.11.2011

4 Detailed

design

21.11.2011 12.01.2012

5 Airfoil

analysis

24.10.2011 31.10.2011

6 Aerodynamic

analysis

21.11.2011 12.01.2012

7 Stability

analysis

21.11.2011 13.01.2012

8 Component

prototyping

20.02.2012 23.03.2012

9 Aircraft

construction

12.03.2012 20.04.2012

III. Conceptual Design

In this chapter we shall talk about the conceptual design investigations held throughout

the first couple of weeks, from the end of October until the middle of November. The initial

design focused on the identification of mission requirements, taking into consideration the

imposed rules and the score awarding system. Many of the proposed design patterns have been

taken out, due to errors of design or unmet technical requirements.

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III.1 Mission requirements

Each aircraft is supposed to meet a number of payload, structural, performance, and

propulsive requirements for the 2012 DBF competition. Each team should design and build a

radio controlled aircraft of limited weight and power, which takes-off in 40 m and flies over the

field to return safely to the runway. The competition is divided in two parts, design and flight:

• In the design part the team will construct a plane considering the requirements and produce a

design report to document the design and construction process as well as financial and teamwork

approach. This design report is reviewed and graded by the competition jury.

• The flight competition will consist of as many runs as possible. The goal is to have at least 3

runs. The number of runs depends on the number of teams and weather conditions. Runs will be

made even under rainy and windy conditions. The decision if it is possible to fly will be made by

the organization team.

•The best performer’s score in each mission normalizes scores for all the other competitors, such

that the best performance receives the maximum allowable points for that mission and other

teams receive a corresponding fraction of the possible points.

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III.2 Translation into design requirements

The aircraft design may be of any configuration except for the rotary wing or lighter-

than- air configurations. The aircraft must take-off only with the energy given by the on-board

propulsion battery pack, and its propulsion is the prescribed electric motor and no form of

externally assisted take-off is allowed. The motor must be an BL Outrunner C4015/1000 and it

must be the only one , the regulations state that the plane must be driven by a single motor and it

must be fitted with a regulator electronic Brushless 30A GX Series. The maximum current is

limited to I = 25A for the competition flights.

The batteries needs have a minimum capacity to ensure the planes can perform at least

one flight pattern. The battery must be LiPo AIRSOFT GENS ACE 11.1V/1600 mA/20C model.

The connector that makes the link between ESC and Battery must be DEANS type connector.

Only one propeller is allowed and the use of a metal one is forbidden. A spinner or

security screw must be also used. The propeller must be a commercial and tested product with

the safety precautions respected.

For the transmission gears, chains and propellers shafts are allowed as long as the

rotation ratio between the motor and the propeller is 1:1.

The maximum take-off weight must not exceed 2000g, leaving the possibility of varieties

of the aircraft design.

No autopilot or control assistance systems may be used. Mixing abilities in the

transmitter/receiver may be used, as long as they do not use any input of sensors.

The aircraft must be able to perform the stability test made during the technical

inspection or before flights during the flight competition. The wing will be supported both-ends

of the wingspan and should not break within a period of 3 seconds. No autopilot and/or control

assistance systems are allowed.

III.3 Considered configuration

While making the configuration of the aircraft, we took into consideration all the design

requirements stated in the FIA DBF Challenge Regulations, that it had to be easy to manufacture

due to the easy design of the components and all the devices that we are allowed to use and all

the requirements that we need to accomplish:

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1. Structural requirements: the system weight is defined as the weight of all components of

the aircraft including the battery, engine, receiver, landing gear, propeller and servo

system weight. The aircraft must pass this test without failure of any type.

2. Take-off requirements: the maximum takeoff distance for each mission is 40m (wheels

off the runway). It is important to note that the field elevation of approximately 80m and

ambient temperatures at the competition site will potentially reduce air density,

depending on temperature and humidity; no forms of externally help are allowed during

take-off.

3. Propulsion: the propeller must be commercially used, tested and safety with precautions

respected.

The most important reason, that convinced us, was the aerodynamic performances. A scoring

analysis was performed to identify the most sensitive variables in the total flight score and assist

in the translation into design requirements. Additionally, the analysis revealed the importance of

matching battery capacity to the number of laps flown and the relative unimportance of absolute

flight speed.

IV. Preliminary Design

During the preliminary design phase, the team focused on analyzing competition rules to

select an aircraft configuration that would maximize competition score. Sensitivity analyses

identified system weight and loading time as key design drivers, with performance in the

efficiency-based delivery mission a secondary factor.

The preliminary design phase focused on fully developing and refining the details of the

design chosen during the conceptual design phase.

The critical aerodynamic design details were determined to be wing area, aspect ratio, and power

requirements at takeoff and cruise. These parameters were optimized using several in-house

MATHCAD and Excel-based performance codes, as well as commercial tools such as XFLR.

Finally, stability, control, and propulsion system analysis over the entire velocity range of

the aircraft was conducted to further refine the design.

Component Types

Wing Monoplane

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Fuselage Convetional

Empennage Convetional

V-tail

H-tail

Landing gear Tricycle

Propulsion tractor

IV.1 Critical design parameters

We realized very quickly, during the brainstorm, that few concepts actually offered

significant advantages in terms of weight, simplicity over a fabric pocket design or a rigid box

design. A rigid design could potentially serve as the primary aircraft structure. The simplicity

and performance per weight of the monoplane would make it the frontrunner. Despite this, the

span and aspect ratio limitation made a multi-wing aircraft an attractive option. A flying wing

configuration was considered for the fact that it would eliminate the fuselage component of the

aircraft. Drawbacks include difficulty of manufacturing and takeoff.

Conventional fuselage is retained for more detailed analysis due to the easy construction

and minimization the time of assembly. The T-tail plane surfaces are kept well out of the airflow

behind the wing, giving smoother flow, more predictable design characteristics, and better pitch

control. As a drawback the aircraft will tend to be much more prone to a dangerous deep stall

condition, where blanking of the airflow over the tail plane and elevators by a stalled wing can

lead to total loss of pitch control. The V-tail is lighter, has less wetted surface area, and thus

produces less drag, but was not considered due to the area required to achieve control equivalent

to a conventional tail resulted in no savings in system weight. The conventional configuration

tail was retained for more detailed analysis; the former for its low risk and the latter for the

possible weight advantage if combined with a reflexed wing airfoil.

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Based on the input regarding the limited take-off length and ground stability, a steerable

tricycle landing gear type was retained for further analysis.

A sample of commonly available electric motors showed a clear trend – the smaller

motors consistently had higher power density. Given the importance of system weight in total

flight score we chose to analyze further the tractor propulsion configuration.

IV.2. Mission Analysis

A scoring analysis was performed to identify the most sensitive variables in the total flight score

and assist in the translation of the above mission requirements into design requirements.

Additionally, analysis revealed the importance of matching battery capacity to the number of

laps flown to complete the mission and the relative unimportance of absolute flight speed. The

mission analysis indicated that a competitive design would minimize system weight and

precisely match the delivery battery weight to the energy needed to complete a chosen number of

laps.

Current Tension Traction Pout Pin

14.3 11.3 873.8 124 168

12.98 10.7 789 107 144

11.7 10 708 91 122

16.6 11.7 870 191

16.4 11.4 860 180

16 11.35 825 177

15.6 11.25 810 173

15.2 11.1 800 168

15 11 795 122.487884 165

15 10.85 780 118.033779 159

14.5 10.75 760 115.064376 155

14.3 10.65 755 111.352622 150

13.5 10.25 715 97.9903071 132

The efficiency of the battery after the analysis was η ≈ 74%

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In order to achieve the predicted drag and lift, each member of the team has chosen a

different airfoil for both wing and empennage that performed well at low Reynolds numbers and

that also had the maximum lift coefficient. The analysis of these factors was performed in

XFLR5 program at the same range of Reynolds numbers i.e. = 80000 = 200000.4

0

200

400

600

800

1000

0 60 120 180 240 270

Tractiune

Tractiune

0

2

4

6

8

10

12

14

16

0 60 120 180 240 270

Curent

Tensiune

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IV.2.1 Wing

After carefully comparing the properties and behavior of every airfoil, the selected airfoil is

NACA 1412.

Graphs of Drag polar

Wing airfoils

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NACA 1412 characteristics:

Thickness: 12.0%

Max CL: 1.098

Camber: 1.0%

Max CL angle: 15.0

Trailing edge angle: 16.5o

Max L/D: 41.806

Lower flatness: 64.9%

Max L/D angle: 6.0

Leading edge radius: 3.5%

Max L/D CL: 0.836

Stall angle: 7.0

Zero-lift angle: -1.0

This airfoil provides an optimum range of Lift over Drag values, allowing the aircraft to

be flown efficiently at a number of cruise attitudes. The NACA 1412 airfoil is a high lift airfoil,

which is required for an aircraft with the design constraints provided in the contest. For

aerodynamic modeling, the lift is optimized by varying wing span and area. The system of

equations used in the optimization program sets the required lift for certain performance

situations, and optimizes the wing dimensions to fit the best design. To provide the lift required

and minimal drag, the wing is designed for an incidence of -4°. This is based on the high lift

airfoil design, an optimum lift to drag ratio, and larger range to advert stall. The calculated lift to

drag ratio was plotted against angle of attack, which is shown in the figure.

Lift Drag Ratio with respect to Angle of attack

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IV.2.2 Empennage (Vertical and horizontal stabilizers) Empennage airfoils :

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The NACA 0008 airfoil was selected for both horizontal and vertical tail due to the symmetric

profile and ease of construction. This airfoil is thin, lightweight, and commonly used for most

tails. The tail was designed to be mounted at an incidence of 0°.

NACA 0008 characteristics:

Thickness: 8.0%

Max CL: 0.692

Camber: 0.0%

Max CL angle: 12.5

Trailing edge angle: 9.7o

Max L/D: 30.528

Lower flatness: 45.6%

Max L/D angle: 4.5

Leading edge radius: 0.8%

Max L/D CL: 0.528

Stall angle: 4.5

Zero-lift angle: 0.0

IV.3. Design and sizing trades

Each team member began by creating a series of models to estimate the weight of wing and

tail surfaces, size tail surfaces based on wing span and tail arm lengths, and relate plan form

limitations to possible aircraft dimensions.

The preliminary aircraft optimization resulted in a basic aircraft geometry which served as a

starting point for the design and refinement of the structure, propulsion system, detailed

aerodynamics, and stability characteristics.

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To select the best aircraft configuration for the contest requirements, assembly methods

for both the wing and tail were analyzed in addition to the traditional tail and wing

configurations.

The primary variable that can be changed to size the aircraft is the wing chord.

IV.3.1 Wing

Having in mind the need to simplify the wing construction, a trapezoidal mono-wing was

chosen. To size the wing, the wingspan was held constant at 500 mm and the chord was varied.

The wing tip chord of 76 mm and a root chord of 130 mm were chosen to optimize the total

flight score.

IV.3.2 Tail

The horizontal tail must provide enough momentum to rotate for takeoff and provide

longitudinal stability. This tail size with a conventional elevator provided the needed momentum

to rotate for takeoff, but an initial stability analysis yielded a high static margin.

Since the tail slides into the fuselage, the root chord of the horizontal tail was limited to

130 mm; therefore, the tail was sized to a 250 mm span by an 76 mm tip chord.

The vertical tail was also sized using previously designed tail volumes. The vertical tail

root chord was sized to 140 mm to stay close to the chord of the horizontal tail, while the tip

chord was 84 mm for aesthetic reasons. This required the overall span of the vertical tail to be

150 mm.

IV.3.3 Fuselage

The selected fuselage starts by having a cylindrical nose with a diameter of 114 mm that

continues with a parallelepiped which is decreasing in volume until reaches the length of 900

mm.

IV.3.4 Propeller

Only single propeller is allowed. However, the use of metal propellers is forbidden. We

had to mention a maximum of three types of propellers in the report which we have chosen to

use in the flight competition. The organizers will provide the team with one GWS

Direct Drive 10 x 6 propeller, which has to be one of the 3 presented in the report.

The chosen propeller provides the adequate amount of thrust for the aircraft’s mission.

Thus larger propellers were not needed to gain the thrust needed for takeoff. Therefore the team

decided to use the one provided by the organizers.

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IV.3.5 Motor

The motor is a BL Outrunner C4015/1000 which has an efficiency and torque enough to drive

propellers. The characteristic of this type of motors are:

maximum intensity of current is 27

Amps;

mass of 90 grams;

maximum traction of 1400 grams;

rotation speed of 1000 rpm;

dimension D 40 x 15 mm.

IV.3.6 Landing gear

A bow-shaped landing gear was chosen for simplicity of construction and mounting.

The width of the landing was designed to ensure stable ground handling, and the height of the

landing gear gives clearance for the large propeller.

IV.4 Stability and control analysis

IV.4.1 Stability analysis

The aircraft we have chosen was optimized to develop a versatile enough design to

handle all mission requirements. With carefully selected wing and tail airfoils, a static

longitudinal stability analysis was able to balance the aircraft while placing both the wing and

empennage incidence angles at 0 degrees from the aircraft fuselage.

The process for achieving static stability for our aircraft configuration required refining a

center of gravity model and updating a stability spreadsheet that provided the necessary

calculations and graphs for both static and dynamic stability.

The major contributors to pitch stability are the wing, tail, and fuselage. Because our

aircraft uses a high lift airfoil, a large negative pitching moment was produced.

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IV.4.2 Control analysis

The aileron roll capability was calculated for mission requirements. The ailerons create a

differential lift that counteract the payload roll moment. The aileron roll equation is unique from

traditional aerodynamic parameters in that it is a function of aileron deflection (β) and velocity.

For future analysis, the aileron moment coefficient was presented within the aileron roll moment

equation below.

This aileron moment equation is based upon strip theory and used to analyze roll control

methods. Because flight velocities are lowest at takeoff and landing, the goal was to create

enough aileron moment to counteract the asymmetric loading before reaching aircraft stall

velocity. After the velocity increases, the roll moment from the ailerons is sufficient to control

the aircraft in flight.

To achieve static and dynamic longitudinal control, the elevator must be properly sized.

In order to balance the aircraft for all necessary trim conditions, the elevator to horizontal tail

area ratio was determined to be 2/3. The size of the elevator attains a 6.13 cm chord and 22 cm

span. This elevator control allows for trim over a spread of 15 degrees angle of attack and was

considered sufficient for all maneuvers required in the mission.

The selected rudder was sized similar to the elevator with a 6.1 cm chord-wise length and 12 cm

in vertical span.

IV.5 Structural analysis

The structure developed met the specified requirements based on the sizing of the

aerodynamic and propulsions calculations performed by the team. Utilizing the selected

conceptual design and benchmarking previous successful aircraft, trade studies were conducted

to select structural configurations.

IV.5.2 Fuselage analysis

After analyzing the configuration and design of the fuselage, we got to the conclusion

that it has a high lift coefficient.

The input data were taken from the stability analysis previously made in XFLR5. Also

some of the data were taken from the actual airplane design measurements like nose, fuselage

and backward surfaces.

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IV.5.3 Landing Gear

To accomodate the range of payload weight distribution for the mission, the landing gear

design is paticularly important. A tricycle configuration with additional outriggers was selected

in the conceptual design phase to provide the ground handling characteristics required for

asymmetrical loads. For the main landing gear, carbon fiber composites were selected for high

strength-to-weight ratio; and the bow shape of the main gear allows for impact absorption on

landing. Carbon fiber composites were also selected for the outriggers. While not a primary

support mechanism, the outriggers must also be strong enough to withstand a hard landing, as it

is anticipated that the plane will land at the end of the mission.

V.Detailed design

Having experience from the conceptual and preliminary design, we had an excellent

starting point for the detailed design phase. After certain brainstorms, we had a clear

understanding of the issues that needed to be complete in this final design phase.

During the detailed design phase, every specific sizing and operational parameter was

settled so that the aircraft could be manufactured effectively and accurately. During this phase

the geometrical properties were decided.

V.1. Dimensional parametres

The dimensional parameters for the final design consist of dimensions for the aircraft, as

seen in the table below:

Fuselage Wing

Lenght (cm) 90 Airfoil NACA 1412

Diameter(cm) 11,4 Span(cm) 500

Hight(cm) 12 Tipchord(cm) 7,6

Tip lenght(cm) 20 Root chord(cm) 13

Tip aspect ratio Aria(cmp)

Backward lenght(cm) 20 Aspect ratio

Aspect ratio Incidence angle(deg)

Section aria(cmp) Aileron area(cmp)

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Horizontal stabilizer Vertical stabilizer

Airfoil Naca0008 Airfoil Naca0008

Span(cm) Span(cm)

Tip chord(cm) Tip chord(cm)

Root chord (cm) Root chord (cm)

Incidence angle(deg) Incidence angle(deg)

Elevator aria(cmp) Rudder aria(cmp)

V.2. Weight and balance

The weight and balance information for the final aircraft design can be seen below. In the

table below, the individual weights for the major components in the aircraft can be seen along

with the various total weights for the aircraft based on all possible payload configurations across

the mission.The table also shows the cg shifts in the lateral and vertical directions.

Weight and balance

Component Weight(g) Xcg(cm) Ycg(cm) Zcg(cm)

Structure Fuselage

Wing

Vertical Tail

Horizontal

Tail

Propulsion Battery

Motor

Propeller

Landing gear Front

Rear

V.3.Flight performance parametres

Using the final aircraft design, flight performance parameters were calculated. The

general aerodynamic and flight performance qualities

for the aircraft are shown in the table below.

Aircraft Parameters

Cl0

Cd0

Clmax

L/D

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V.4. Mission Performance

After the design was finished, mission performance was reevaluated for more

accurate goals and expectations for mission performance.

Mission Performance

Cruise speed

Maximum speed

Takeoff Distance

Stall incidece

V.5. Drawing phase

In this phase, the plane was entirely drawn using SolidWorks. The following pages

consist of the following drawings: Aircraft 3-View, Systems Layout;

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VI. MANUFACTURING PLAN AND PROCESSES

This phase of the project centered on planning , executing the creation of components and the

assembly of the aircraft itself . A great deal of planning was done in scheduling, construction techniques

and materials used , due to the time restrictions . The manufacturing plan was divided into four major

components: Fuselage, Wings, Tail, Landing Gears. The following manufacturing plan defines the

alternatives investigated, and the processes selected.

A set of figures were established to help in the manufacturing decision making process for the

manufacturing of all the aircraft components . The construction method must provide enough structural

integrity to the system so that it can perform all the required missions and it must have a relatively small

time of implementing , because of this we have chosen four figures : ease of construction, structural

integrity maintainability and weight.

VI.1 Investigation and selection of major components and assemblies

Many manufacturing processes were researched to determine the most reliable light weight

process for each major assembly though we listed in the following section only the most appealing ones

for every major component.

VI.1.1 Fuselage

The fuselage had major methods considered for its construction, however we chose the XPS

material leaving only one method available. While weight and construction time were still the dominating

factors for this piece of the aircraft, maintenance was also highly considered due to its integration role for

all the other components of the aircraft.

•Extruded polystyrene foam (XPS) consists of closed cells, offers improved surface roughness and higher

stiffness and reduced thermal conductivity. The density range is about 28 – 45 kg/m3.

Extruded polystyrene material is also used in crafts and model building, in particular architectural models.

Because of the extrusion manufacturing process, XPS does not require facers to maintain its thermal or

physical property performance. Thus, it makes a more uniform substitute for corrugated cardboard.

Thermal resistivity is usually about 35 m·K/W (or R-5 per inch in American customary units).

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VI.1.2 Landing Gear

The landing gear consisted of three different pieces: main gear, rear wheel, and outriggers.

Benchmarking previous designs, a composite layup construction method was used. The molds for the

main gear and rear wheel are going to be made of high density foam which will provide a smooth surface

for composite layup. The final construction layup of the main gear consisted of a carbon fiber composite

layup. The rear gear will be also a composite layup, but with a Kevlar fabric core, which will provided

durability and a higher spring constant.

VII. References

XFLR5, 24 Oct, 2011. <http://xflr5.sourceforge.net/xflr5.htm>

UIUC Airfoil Coordinate Database.” 24 Oct, 2011.

SolidWorks 2010

Mihai M.Nita,Florentin V.Moraru si Radu N.Patraulea

„Avioane si rachete – concepte de proiectare”, Ed. Militara Bucuresti 1985

Wikipedia