Airplane report
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Transcript of Airplane report
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