ENGI 7926: Detailed Design Report
ENGI 7926: Detailed Design Report
Andrew Batstone (200818938)
Evan Macleod (200722171)
Nick Robbins (200716900)
John Weir (200703957)
Jad Yassine (200856078)
7/4/2012
Submitted To: Dr. James Yang
ENGI 7926: Detailed Design Report
Table of Contents
Table of Appendices ....................................................................................................................................... i
Table of Figures ............................................................................................................................................. ii
Table of Tables ............................................................................................................................................. iv
Table of Equations ........................................................................................................................................ v
List of Acronyms and Abbreviations ............................................................................................................ vi
Executive Summary ..................................................................................................................................... vii
1.
Design Overview ................................................................................................................................... 1
1.1.
Fuselage ........................................................................................................................................ 1
1.2.
Wing .............................................................................................................................................. 1
1.2.1.
Airfoil ..................................................................................................................................... 1
1.2.2.
Airfoil Concepts Comparison ................................................................................................ 1
1.2.3.
Wing Control Surfaces........................................................................................................... 2
1.2.4.
Wing Sizing ............................................................................................................................ 2
1.3.
2.
3.
Tailplane ........................................................................................................................................ 2
1.3.1.
Tailplane Design .................................................................................................................... 2
1.3.2.
Tailplane Control Surfaces .................................................................................................... 4
1.4.
Propeller Testing ........................................................................................................................... 5
1.5.
Landing Gear and Tailwheel .......................................................................................................... 5
Structural Analysis................................................................................................................................. 5
2.1.
Fuselage Structural Analysis ......................................................................................................... 5
2.2.
Wing Structural Analysis ............................................................................................................... 6
2.3.
Tailplane Structural Analysis ......................................................................................................... 6
2.4.
Landing Gear Structural Analysis .................................................................................................. 6
2.5.
Assembly Structural Analysis ........................................................................................................ 7
Aerodynamic Analysis ........................................................................................................................... 7
3.1.
Fuselage Aerodynamic Analysis .................................................................................................... 7
3.2.
Wing Aerodynamic Analysis .......................................................................................................... 8
3.3.
Tailplane Aerodynamic Analysis ................................................................................................. 10
3.4.
Assembly Aerodynamic Analysis ................................................................................................. 10
ENGI 7926: Detailed Design Report
4.
Vibration Analysis ............................................................................................................................... 12
5.
Stability Analysis ................................................................................................................................. 14
5.1.
CG Determination ....................................................................................................................... 14
5.2.
Euler Stability Derivatives ........................................................................................................... 14
6.
Constraints Criterion Checklist............................................................................................................ 15
7.
Objectives List ..................................................................................................................................... 16
8.
Justification of Design Changes........................................................................................................... 17
9.
Project Management Plan .................................................................................................................. 17
10.
Conclusion ....................................................................................................................................... 18
11.
References ...................................................................................................................................... 19
12.
Acknowledgements ......................................................................................................................... 19
ENGI 7926: Detailed Design Report
Table of Appendices
Appendix A - Foil Shapes
Appendix B - Balsa Tensile Stress Testing
Appendix C - SAE Carrying Case Specifications Check
Appendix D - Drop Test Criteria
Appendix E - Control Surface Servo Torque
Appendix F - Engine Cover Heat Transfer Study
Appendix G - Structural Analysis
Appendix H - Project Management Plan
Appendix I - Technical Drawing Package
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ENGI 7926: Detailed Design Report
Table of Figures
Figure 1 - Fuselage Aerodynamic Analysis in SolidWorks ............................................................................. 8
Figure 2 - XFoil Lift/Drag Comparison ........................................................................................................... 8
Figure 3 - Center of Pressure and Induced Angle ......................................................................................... 9
Figure 4 - Wing Simulation in Xfoil ................................................................................................................ 9
Figure 5 - Flow Trajectory Analysis for Tailplane ........................................................................................ 10
Figure 6 - Flow Trajectory Analysis for Assembly ....................................................................................... 11
Figure 7 - Lift and Drag Forces on Assembly ............................................................................................... 11
Figure 8 - Example Harmonic Response...................................................................................................... 13
Figure 9 - Amplitude Ratio Plot ................................................................................................................... 13
Figure 10 - Center of Gravity ....................................................................................................................... 14
Figure 11 - Dynamic Stability Eigenvalue Plot ............................................................................................. 15
Figure 12 - Project Management Plan Legend ............................................................................................ 18
Figure 13 - Selig 1223 and Eppler 420 Airfoils ........................................................................ Appendix A - 1
Figure 14 - UI-1720 and CH-10 Airfoils.................................................................................... Appendix A - 1
Figure 15 - Stress vs. Load of Balsa Wood ...............................................................................Appendix B - 1
Figure 16 - Balsa Testing Apparatus .........................................................................................Appendix B - 1
Figure 17 - Balsa Testing Apparatus .........................................................................................Appendix B - 2
Figure 18 - Balsa Testing Apparatus .........................................................................................Appendix B - 2
Figure 19 - Balsa Fracture ........................................................................................................Appendix B - 3
Figure 20 - SAE Carrying Case Sizing Check ..............................................................................Appendix C - 1
Figure 21 - Heat Transfer Testing Results ................................................................................ Appendix F - 1
Figure 22 - Engine Cover Material Specifications .................................................................... Appendix F - 1
Figure 23 - Isoclipping from Heat Transfer Test ....................................................................... Appendix F - 2
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ENGI 7926: Detailed Design Report
Figure 24 - Heat Testing Results............................................................................................... Appendix F - 3
Figure 25 - Wing Structural Analysis ....................................................................................... Appendix G - 1
Figure 26 - Displacement and Von Mises Stress on Horizontal Stabilizer .............................. Appendix G - 1
Figure 27 - Displacement and Von Mises Stress on Vertical Stabilizer ................................... Appendix G - 2
Figure 28 - Landing Gear Stress Analysis, 52.0375 N Force .................................................... Appendix G - 2
Figure 29 - Landing Gear Displacement Analysis, 52.0375 N Force ........................................ Appendix G - 3
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ENGI 7926: Detailed Design Report
Table of Tables
Table 1 - Airfoil Concepts Comparison .......................................................................................................... 1
Table 2 - Tailplane Area Analysis................................................................................................................... 3
Table 3 - Propeller Testing ............................................................................................................................ 5
Table 4 - Inputs for SolidWorks Fuselage Flow Simulation ........................................................................... 7
Table 5 - Fuselage Aerodynamic Forces ........................................................................................................ 7
Table 6 - Dynamic Stability Parameters ...................................................................................................... 15
Table 7 - Constraints Criterion Checklist ..................................................................................................... 16
Table 8 - Objectives List .............................................................................................................................. 17
Table 9 - Nose Drop Test Details ............................................................................................. Appendix D - 1
Table 10 - Bottom Drop Test Details ....................................................................................... Appendix D - 1
Table 11 - Servo Torque Requirements ................................................................................... Appendix E - 1
Table 12 - Heat Transfer Testing Parameters .......................................................................... Appendix F - 1
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ENGI 7926: Detailed Design Report
Table of Equations
Equation 1 - Stall Speed ................................................................................................................................ 2
Equation 2 - Tailplane Range of Motion ....................................................................................................... 4
Equation 3 - Maximum Control Surface Force .............................................................................................. 4
Equation 4 - Maximum Control Surface Moment......................................................................................... 5
Equation 5 - Landing Gear Force Calculation ............................................................................................... 6
Equation 6 - Frequency of Motor................................................................................................................ 12
Equation 7 - Stiffness of Balsa Specimen (Motor Mount Assembly) .......................................................... 12
Equation 8 - Natural Frequency of Balsa Specimen (Motor Mount Assembly) .......................................... 12
Equation 9 - Frequency Ratio ...................................................................................................................... 12
Equation 10 - Amplitude Ratio (Motor Mount Assembly) .......................................................................... 12
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ENGI 7926: Detailed Design Report
List of Acronyms and Abbreviations
A: Current (Amps)
Aero (abbrev.): Aeroplane
CG: Center of Gravity
CL: Center of Lift
ESC: Electronic Speed Controller
FDM: Fused Deposition Modeling
FEA: Finite Element Analysis
ft/s: Feet per second (Imperial speed measurement)
g: Gravitational constant (9.81 m/s2)
in (abbrev.): Inch (imperial length measurement)
kg (abbrev.): kilogram (metric mass measurement)
λ: Taper ratio
lb: pound (imperial mass measurement)
mm (abbrev.): Millimetre (metric length measurement)
MUN: Memorial University of Newfoundland
m/s: meters per second (metric speed measurement)
N: Newton (metric weight measurement)
Prop (abbrev.): Propeller
psf: pounds per square foot (imperial pressure measurement)
RC: Remote Controlled
SAE: Society of Automotive Engineers
Servo (abbrev): Servomotor
W: Power (Watts)
V: Voltage (Volts)
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ENGI 7926: Detailed Design Report
Executive Summary
MUNA Enterprises has successfully completed the design phase of ENGI 7928: Mechanical Engineering
Design Project. By using theory and concept generation and selection, a definitive remote controlled
aeroplane concept was created. The design was refined based on a number of concept screening and
scoring matrices carried out for various plane sections; these sections included the wing, fuselage and
tailplane design of the plane. It was also necessary to conduct propeller testing, structural finite element
analysis, and crash/impact finite element analysis during the design phase. Additionally, a multitude of
stability analysis was conducted, including computer modeling using AVL.
Although most components are being designed by MUNA Enterprises, some material purchasing was
also required during the design phase of the project. The materials and assemblies purchased thus far
include: landing gear assembly, tail wheel assembly, build material (balsa and Monokote), engineering
adhesives and adhesive applicators, Monokote heat gun, fasteners (various screws, etc.), and servo
pushrod and control horn material.
With a total wingspan of 1205.2 mm, a tip-to-tail length of 843.7 mm, an estimated lift force of 35 N and
a maximum total loaded weight of 8.07 pounds, the estimated target payload fraction for MUNA
Enterprises' "Balsa Salsa" is .743.
The aircraft design features several innovations; most notably the use of fabricated interlocking bars in
combination with a high-wing mount that mimics foil shape for rapid wing fastening. Wings features
include an Eppler 420 airfoil, skeletal rib work for maximum material reduction, and a traditional aileron
control surface. The fuselage innovations consist of an ABS-M30 nosepiece capable of withstanding
crash impact, accurately housing the motor, and ventilating the motor to optimal thermal performance
conditions. A shelving unit has also been placed within the fuselage for quick and easy access for
payload customization. The tailplane features a fast-assembly boom mount and a lightweight, minimaldrag design.
Stability analysis via traditional RC aero design principles and AVL shows that the “Balsa Salsa” is stable
when considered in both static and dynamic conditions. The plane has been determined to have
desirable maneuverability characteristics due to the fact that the CG falls between the leading edge of
the wing and the CL.
The project is on schedule, and the initial fabrication of "Balsa Salsa" is projected to be completed as of
July 7th, 2012. The testing will commence as of July 7th, 2012. If necessary, redesign will be executed
between July 9th and July 15th. The project cost to date is $192.19, well within budget constraints
assigned by Dr. James Yang.
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ENGI 7926: Detailed Design Report
1. Design Overview
1.1.
Fuselage
The fuselage designed by MUNA Enterprise for their RC Aero "Balsa Salsa" provides a well balanced
concept that considers aerodynamics, strength, machinability and overall weight. The fuselage originally
consisted of a box-shaped structure that would provide ease for machinability and assembly, however
the drag force and weight were too high for MUNA Enterprises' standards. Numerous iterations were
made to the original design such as tapering the body of the plane to create an improved streamline as
well as cutting a truss structure for the side walls to reduce weight. Furthermore, a highly aerodynamic
nose has been designed and will be manufactured with MUN's FDM machine from Fortus ABS-M30
material. With all sides angled to improve the streamline of the fuselage, it was determined that
machinability would be too challenging and assembling all parts would take a high degree of precision.
Therefore, the strengths of both concepts were combined to result in the final fuselage model. For
further information on design and assembly, please see Appendix I - Technical Drawing Package.
1.2.
Wing
1.2.1. Airfoil
One of the most important steps in the design of an aircraft is the selection of airfoil. Much of the
conceptual design focused around choosing an airfoil with optimal properties to provide maximum lift at
low speeds, while minimizing drag and being easy to design around and fabricate. Through much
research, MUNA Enterprises has chosen four airfoils based on their high coefficients of lift and low drag.
These foils are the Selig 1223, Eppler 420, UI-1720 and CH-10. The general shapes of these foils can be
seen in Appendix A - Foil Shapes.
1.2.2. Airfoil Concepts Comparison
Airfoil
Selig 1223
Eppler 420
UI - 1720
CH-10
Lift
4
3
1
2
Ease of Construction
1
4
2
3
Drag
1
2
4
3
Sum
6
9
7
8
Rank
4
1
3
2
Table 1 - Airfoil Concepts Comparison
In evaluating the four different airfoil concepts, three categories in which the airfoils could be measured
were considered. These were lift, ease of construction, and drag. Lift and drag values were calculated
using Xfoil. The airfoils were then rated from one to four with four being the best rating. The scores
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ENGI 7926: Detailed Design Report
were then totalled and the airfoil with the most points was selected. According to the matrix, the Eppler
420 foil was chosen because of its combination of flight characteristics and ease of construction. More
data on lift and drag can be seen in section 3.2. Detailed drawings of the wing can be seen in Appendix I
- Technical Drawing Package.
1.2.3. Wing Control Surfaces
The roll of the plane is controlled by the wing ailerons. Ailerons are sized to be 10-20% of the wing area,
or 50% of the wing semi span and 25% of the wing chord. The ailerons on MUNA Enterprises’ plane are
approximately 230mm x 66mm. These values are relatively close to the 50% wing semispan and 25%
wing chord that is recommended. This creates a 0.03036 m2 total aileron area, which is 15% of the wing
area. This also meets additional recommended criteria. The required servo torque for the ailerons can
be found in Appendix E - Control Surface Servo Torque.
1.2.4. Wing Sizing
The wings were sized as to generate the most lift possible, however the wings’ main constraint is the
fact that they must fit in the SAE-specified foam fitting carrying case. The case dimensions being
24"x18"x8" meant that each wing must be less than 2 feet long. Knowing this, MUNA Enterprises chose
to employ a 500mm wing length to leave an adequate amount of room in the case for additional
components. The cord length was calculated using the stall speed equation, as seen in Equation 1
where:
Vstall is the stall speed in m/s
W is the weight of the loaded plane in kg
g is gravity in m/s2
 is the density of air at flight conditions in kg/m3
S is wing area (Wing Span * Cord Length) in m2
and CLmax is the maximum wing lift coefficient
Equation 1 - Stall Speed
Vstall 
Wg
1
  S  C L max
2
Using this equation, and knowing all variables except V and Cord Length, and assuming a stall velocity of
half the static speed, or 12m/s, gives a cord length of approximately 200mm.
1.3.
Tailplane
1.3.1. Tailplane Design
The tailplane design which MUNA Enterprises used is a conventional-style configuration, featuring a
controllable rudder and elevator. This follows the design selected following the concept screening and
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ENGI 7926: Detailed Design Report
scoring matrices for tailplane design. Both the vertical and horizontal stabilizers are designed using a
3.18 mm thickness (1/8 in) balsa frame with Monokote wrap surface coating. Weight was successfully
reduced for these components without sacrificing structural integrity by using truss framework. To
minimize the weight and enhance structural strength, no airfoil has been modeled for either the
horizontal or vertical stabilizer. As a viable substitute, the 3.18 mm thick balsa sheets coated with
Monokote will essentially serve as the foil shape for both stabilizers. It was determined that the amount
of lift generated by a tailplane crafted from an airfoil similar to the wing would be negligible, therefore
the balsa framework was selected for the tailplane design. Additionally, modeling the tailplane with
airfoils would be less structurally sound than using one continuous sheet of balsa, and more complicated
to fix in the event that redesign is necessary. Memorial University's laser cutter has the capability to
accurately and quickly cut materials such as balsa; as such, the laser cutter will be utilized whenever
possible.
Current RC model building practices suggest guidelines for tailplane effective areas; these guidelines
state that in order for vertical and horizontal stabilizers to be effective, they must comprise a certain
percentage range of the aircraft's wing area. For the vertical stabilizer and rudder, common practice
shows that areas between the ranges of 25-30% will yield control surfaces with negligible slop and high
manoeuvrability. The range that the rudder should comprise of this total is 15-20%. For the horizontal
stabilizer, a range of 30-40% of the wing area will optimize manoeuvrability. A portion of 10-15% of this
total should be dedicated to the elevator according to practice. Table 2 shows the area analysis of
MUNA Enterprises' tailplane design.
Description of Area
Vertical Stabilizer
Rudder
Vertical Stabilizer + Rudder
Horizontal Stabilizer
Elevator
Horizontal Stabilizer + Elevator
Area
(mm2)
24193.5
5806.4
29999.9
28129.0
11225.8
39354.8
% of Total
Stabilizer Area
80.7
19.4
100
71.5
28.5
100
% of Wing
Area
22.2
5.3
27.5
25.8
10.3
36.1
Table 2 - Tailplane Area Analysis
The areas required for airplane acrobatic manoeuvres are substantially larger than those shown in Table
2, however acrobatics is not a requirement of SAE`s Micro Class planes. Although MUNA Enterprises
predicts that using a tailplane of these specifications will enhance the plane’s manoeuvrability, both
pitch and heave control are of greatest importance considering the flight path a straight line. The
elevator will mainly be used to mitigate the heave and pitch variability during flight path (i.e. “porpoise”
effect), and its secondary use will be the addition of a marginal amount of lift. Detailed drawings of the
tailplane can be found in Appendix I - Technical Drawing Package.
The entire tailplane assembly is capable of fitting in the SAE Micro-Class regulated carrying case,
therefore disassembly of the tailplane is not required. This can be observed in drawing Appendix C - SAE
Carrying Case Specifications Check.
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ENGI 7926: Detailed Design Report
1.3.2. Tailplane Control Surfaces
The tailplane features two conventional control surfaces: a rudder, for yaw and sway control, along with
an elevator for heave and pitch control. One servo is required for control of each surface for a total of
two servos. The servos have been placed on the tailplane itself to eliminate use of a long pushrod;
placing the servo close to its control surface allows for greater reliability, ease of access, and less chance
of pushrod buckling while having a marginal effect on the airplane’s CG. The clevis pins used to hold the
servo pushrods were purchased from Signal Hobbies. They are designed to be lightweight and low-drag,
and they offer sufficient geometric dimensions necessary to produce a moment arm for each control
surface. The range of motion for each control surface was calculated based on the translational motion
which can be provided from each servo. Equation 2 gives the servo’s effective circumference, which can
be used to determine the range of motion (Θrange) for each control surface:
Equation 2 - Tailplane Range of Motion
Cservo = 0.5*π*rservo
ΘControlSurface = sin-1(Cservo / LControlSurface
Θrange = 2*ΘControlSurface
Where:
Cservo is the quarter-circumference of the servo rotation,
rservo is the servo radius,
ΘControlSurface is one half of the angle which the control surface can be changed by (i.e. CW or CCW
rotation),
and LControlSurface is the length of the control surface.
Once ΘControlSurface and a flight speed (Vflight) are known, a subsequent maximum wind loading force (Fmax)
and moment (Mmax) can be found as follows:
Equation 3 - Maximum Control Surface Force
Fmax = A*P*Cd
Where:
A is the projected surface area of the object (ft2),
P is the wind pressure (psf), approximated as 0.00256*Vflight2,
and Cd is the drag coefficient (2.0 for flat plates).
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ENGI 7926: Detailed Design Report
The maximum moment is calculated using the following equation:
Equation 4 - Maximum Control Surface Moment
Mmax = Fmax*LCG
Where LCG denotes the perpendicular length from the center of the distributed wind load to the plane`s
CG.
1.4.
Propeller Testing
Two propellers were tested using MUN's propeller testing apparatus. The apparatus measured the
thrust and current draw of the propellers using the receiver, battery, ESC and transmitter to control the
motor. The results of the testing can be seen in Table 3.
Prop Size
Current Draw
Thrust
(A)
(kg)
11X5.5
17
1.2
12X6
23
1.5
(diameter X pitch in inches)
Table 3 - Propeller Testing
Knowing the provided ESC has a maximum current of 35Amps, it can be seen that either of these
propellers can be used. MUNA Enterprises has decided on using a 12 in X 6 in propeller as it has an
adequate safety factor on the current draw, and develops more thrust and speed than the 11 in X 5.5 in
propeller.
1.5.
Landing Gear and Tailwheel
Both the landing gear and tailwheel were sized using the relative dimensions of the assembly, and
purchased via Hobby King. Additional details are available via bills of materials in Appendix I - Technical
Drawing Package.
2. Structural Analysis
2.1.
Fuselage Structural Analysis
Two drop simulations have been performed on the fuselage model through SolidWorks Simulation. The
results for both tests were obtained when simulating a drop from a height of approximately 11 m.
However, MUNA Enterprise has concluded that maintaining a low weight is more valuable than
increasing structural support of the RC plane; therefore, the nose has been designed to be as light as
possible. In the event of a drop or crash landing, due to the ease of fabrication of the fuselage
components, repair is quick and inexpensive. Additionally, MUNA Enterprises is confident that the
manoeuvrability of the "Balsa Salsa" will ensure the plane will not experience a significant drop during
the MUN flight competition. The criteria used for the drop test can be seen in Appendix D - Drop Test
Criteria.
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ENGI 7926: Detailed Design Report
2.2.
Wing Structural Analysis
Wing structural analysis was performed with a distributed load of 20 N simulating, a lift of 4.5 pounds
per wing. This resulted with a maximum stress of less than 6 MPa, far below the conservative value for
yield strength of 22 MPa of balsa. The deflection in the wing reaches a maximum of 1.5 mm, well within
the acceptable limits for this application. Figure 25 shows these results and can be seen in Appendix G Structural Analysis.
2.3.
Tailplane Structural Analysis
The tailplane is secured to a balsa tailplane boom via adhesive applied at the vertical stabilizer,
horizontal stabilizer and the boom interfaces. The vertical stabilizer can be secured by interlocking it
with a small slit in the middle of the horizontal stabilizer. The horizontal stabilizer and tailplane boom
interface forms a strong attachment capable of withstanding the flexure, tensile and compressive forces
of balsa (taken as 5.3, 7-22, and 12.1 MPa, respectively). Both stabilizers were analyzed using a typical
loading scenario, with a drag force at 45o angle and a 15 m/s flying speed. A sample of the FEA testing
for the tailplane stabilizers can be seen in Figure 26 and Figure 27 located in Appendix G - Structural
Analysis.
The horizontal stabilizer experiences a maximum loading case deflection of .001 mm, and a flexure
stress of 0.03 MPa. The vertical stabilizer experiences a similar loading scenario due to its similar
geometry; the maximum loading case deflection of the rudder is .001 mm, and the maximum flexure
stress is 0.03 MPa. These values are well within the safety margin for balsa and the nylon hinges used to
connect the control surfaces and the main section of each stabilizer.
2.4.
Landing Gear Structural Analysis
The landing gear assembly is a stock part purchased from Hobby King. The main frame is manufactured
from carbon fiber, making the entire assembly extremely strong while maintaining its lightweight for a
total mass of 22 grams. Below is a structural analysis of the landing gear assembly. The wheels were
removed, and a uniform radial load was placed on both axles. It was determined that the axle locations
are where the critical stresses in the assembly would lie based on theoretical analysis. Using the plane's
total cargo-loaded weight as the landing scenario and applying a safety factor of 1.5, a total load of
52.0375 N was determined using the following equation:
Equation 5 - Landing Gear Force Calculation
Flg = Wloaded plane * SF
Where:
Flg = The force on the landing gear (3.536 kg*g)
SF = The design safety factor (1.5)
One can observe from Figure 28 and Figure 29 in Appendix G - Structural Analysis that the landing gear is
structurally sound, exhibiting a maximum of 0.113 MPa of stress and a maximum deflection of 0.1056
mm. With a safety factor applied to a probably loading case, the landing gear assembly only exhibits
.018% of its ultimate tensile strength, making it strong enough to withstand the force of most severe
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ENGI 7926: Detailed Design Report
crash landing cases. The most likely scenario for this loading case results in the landing gear keeping
intact, while the balsa base of the fuselage will most likely fracture. Due to its ease of fabrication and
capability of interchanging, the fuselage can be easily repaired.
2.5.
Assembly Structural Analysis
Structural analysis was completed on every major component of the airplane separately. This testing
included all major stresses exhibited on each individual part of the plane. MUNA Enterprises believes
this testing adequate to show the "Balsa Salsa" will not structurally fail in standard flight. Additionally,
performing individual component structural analysis provided more accurate loading and fixture
geometry, resulting in greater accuracy of stress and deflection calculations.
3. Aerodynamic Analysis
3.1.
Fuselage Aerodynamic Analysis
The aerodynamic analysis of the final fuselage design was performed using SolidWorks Flow Simulation.
The input values for this simulation can be seen in Table 4.
Parameter
Value
Air Pressure
101.235 KPa
Air Temperature
293.2 Kelvin
Surface Roughness
15 micrometers
Velocity (x-direction)1
15 m/s
Note:
1) X-axis is along fuselage, starting at the nose and ending at the tail.
Table 4 - Inputs for SolidWorks Fuselage Flow Simulation
The results obtained from this simulation were reasonable and show desirable aerodynamics. MUNA
Enterprises believes this fuselage configuration will be well-suited for competition and meets all
constraints previously specified. Table 5 shows numerical values for drag and lift of the fuselage.
Parameter
Value (N)
Drag
0.7146
Lift
-0.1226
Table 5 - Fuselage Aerodynamic Forces
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ENGI 7926: Detailed Design Report
Figure 1 - Fuselage Aerodynamic Analysis in SolidWorks
3.2.
Wing Aerodynamic Analysis
Initial analysis involved calculating lift and drag values for four different foils using the program XFoil.
Figure 2 shows the lift to drag ratio of four foils with respect to a varying attack angle.
Figure 2 - XFoil Lift/Drag Comparison
It can be seen that the Selig 51223 has the best lift to drag ratio, however due to the low machinability
of this foil, an Eppler 420 foil was chosen instead.
After finalizing the selection of the foil, the wing size and the stabilizer size of the plane was modeled in
Xfoil to determine the center of lift and the total lift generated at a set speed of 15m/s. Xfoil was also
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ENGI 7926: Detailed Design Report
used to determine the wing's angle of attack of 5 degrees. It was found that the total lift developed by
the wing in this configuration is approximately 36N.
Figure 3 shows the center of pressure and induced angle of the wing. The center of pressure is the
localized point at which the lift acts. This point is used to calculate the pitch stability of the plane. The
induced angle of attack shows the change of angle of the airflow over the wing in flight, this is due to the
airfoils and wingtip vortices producing a "downwash" that further deflects the local airflow in the vicinity
of the wing downward.
Figure 3 - Center of Pressure and Induced Angle
Figure 4 - Wing Simulation in Xfoil
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ENGI 7926: Detailed Design Report
3.3.
Tailplane Aerodynamic Analysis
The tailplane is designed with conventional rudder and stabilizer control surfaces manufactured from
1/8" balsa, which results in minimized drag forces for the entire tailplane assembly. A flow trajectory
analysis was conducted using SolidWorks at a flying speed of 15 m/s to gauge the presence of drag
forces on the tailplane. An example of the flow trajectory analysis can be seen in Figure 5.
Figure 5 - Flow Trajectory Analysis for Tailplane
As expected, the largest drag forces exhibited by the tailplane result from the boom-stabilizer interface,
where both servomotors are also present. However, the drag forces exerted on the tailplane by both
servomotors are negligible compared to the amount of drag force which both the rudder and elevator
can apply; the servo drag profile ratio is 1.21% of the maximum elevator lift/drag, and 2.34% of the
maximum rudder lift/drag.
3.4.
Assembly Aerodynamic Analysis
The assembly aerodynamic analysis was performed using SolidWorks Flow Simulation. The aircraft
assembly was created by removing all trusses and forming solid bodies in order to correctly calculate
drag and lift forces on bodies. By defining a constant velocity of 15 m/s in the X direction and
establishing a computational domain large enough to suitably model all 3 dimensions of the assembly,
the maximum lift and drag forces can be found for all components of the assembly. The flow trajectory
of the fuselage, high wing mounting system and tailplane can be seen in Figure 6, along with a plot of
the 3-dimensional drag forces in Figure 7.
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ENGI 7926: Detailed Design Report
Figure 6 - Flow Trajectory Analysis for Assembly
Lift Force
Drag Force
Figure 7 - Lift and Drag Forces on Assembly
Figure 7 shows that the X and Y components of force stabilize at 4N and 6 N, respectively. This shows
that the lift is calculated to be 6.43 N and the drag is 4.05 N at a 15 m/s speed. MUNA Enterprises
predicts that the total lift that the "Balsa Salsa" will be able to generate will be approximately 35 N; the
above figures merely represent that the "Balsa Salsa" generates positive lift in its stagnant position at 15
m/s. This model ensures that the forces are constant based on the stabilization of the values for force,
given the total number of 377 iterations.
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ENGI 7926: Detailed Design Report
4. Vibration Analysis
A common problem which can cause catastrophic failure of planes and automobiles is mechanical
vibration. If an engine or piece of rotating equipment is running at the same frequency as the natural
frequency of the structural material, resonance occurs. Resonance causes system instability due to an
amplification of stored vibrational energy, and this energy will increase the amplitude (i.e. displacement)
of a structural material. For MUNA Enterprises' aircraft, it is possible to determine whether the system
will be at resonance during flight speed using the following equations:
ωmotor = 8800 rpm = 921.53 rad/s
Equation 6 - Frequency of Motor
kbalsa =
=
= 2.7458 GN
Equation 7 - Stiffness of Balsa Specimen (Motor Mount Assembly)
ωnatural =
=
= 550152.3 rad/s
Equation 8 - Natural Frequency of Balsa Specimen (Motor Mount Assembly)
r=
=
= 0.0016750
Equation 9 - Frequency Ratio
=
Equation 10 - Amplitude Ratio (Motor Mount Assembly)
Because the frequency of the input force is less than the natural frequency of the system, the harmonic
response of the system, xp(t), is in phase with the forcing function, as shown in Figure 8 and Figure 9.
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ENGI 7926: Detailed Design Report
Figure 8 - Example Harmonic Response
Figure 9 - Amplitude Ratio Plot
Due to the large difference in magnitude between ωmotor and ωnatural, the risk of operating the motor
within close proximity of the natural frequency of the balsa motor mount is negligible. It would be
virtually impossible to observe adverse effects from harmonic resonance within the system.
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ENGI 7926: Detailed Design Report
5. Stability Analysis
5.1.
CG Determination
The CG of the full assembly was found using SolidWorks. The assembly included all electronics such as
battery, ESC, receiver, and motor, as well as the payload. Figure 10 shows the location of the CG, as well
as the CL developed by the wings. It can be seen that the CG lies 63.25 mm closer to the front of the
plane than the CL. This fits the RC model design practice that the center of gravity should lie between the
front of the wing cord and the CL.
Figure 10 - Center of Gravity
5.2.
Euler Stability Derivatives
Measurements and data from the full SoildWorks assembly were used to develop a model for AVL. The
AVL model was used to calculate the dynamic stability of the plane in flight. The dynamic stability
Eigenvalue plot from AVL can be seen in Figure 11, while Table 6 shows this information in a more
readable format.
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ENGI 7926: Detailed Design Report
Figure 11 - Dynamic Stability Eigenvalue Plot
Aircraft Mode Eigenvalues
Mode
Short Period
Phugoid
Roll
Dutch Roll
Spiral
η
ω
ζ
t1/2
-1.84E-02 1.149013 0.016027 37.64015
-2.88399 4.797397 0.601156 0.240344
-17.9675
0 0.038578
-0.82784 5.939283 0.139384 0.837296
-0.12245
0 5.660675
Table 6 - Dynamic Stability Parameters
Table 6 shows that the aircraft is stable in all modes of flight. The Short Period root is very small, yielding
a relatively long time to half amplitude. Short Period is not a significant issue and is easily corrected with
control input from the pilot.
6. Constraints Criterion Checklist
The following constraints checklist has been updated to reflect progress which has been made. The
original constraints checklist was drafted in May 2012. Statuses which are marked with an "M" indicate
that the constraint will be met by the current design. Statuses which are marked with an "R" indicate
that further refinement may be necessary in order for the aircraft to meet the constraint. Statuses which
have changed from "R" to "M" (i.e. progress) have been noted in green text, whereas statuses which
have changed from "M" to "R" will be noted in red text (i.e. steps backward).
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ENGI 7926: Detailed Design Report
CONSTRAINT DESCRIPTION
JUNE 1
STATUS
JUNE 29
STATUS
R
R
M
R/M
M
M
1. Can the aircraft fly in a 20 m straight line in
NL outdoor conditions?
2. Can the aircraft be hand-launched?
3. Can the aircraft carry the maximum
payload possible:
A) Can the cargo bay support a payload 2x2x5
A)
2x2x5
inches in dimension?
inches
in dimension?
B) Will the payload be supported as a
B)
Supported
and retained in a specially
homogenous
mass?
designed cargo bay?
R
M
R
M
R
M
C) Does the payload have 0 lead content?
C) Can the payload be supported as a
homogenous mass?
4. Is the aircraft designed to achieve the lowest
possible
weight (∴
highest
payload
D) Does empty
the payload
have
a lead
content of
fraction
possible)?
0?
R
M
R
M
R
R
R
M
5. Can all operable contents be packaged in a
form-fitting foam carrying case with inside
dimensions of the case less than or equal to
24x12x8 inches?
R
M
6. Can the aircraft be assembled and launched
in less than 3 minutes by 2 people?
R
R
7. Is the aircraft designed so less than 10 in3
material is used in MUN’s FDM machine?
M
M
8. Is the aircraft design heavier than air?
M
M
9. Are non-metal propellers being used for the
aircraft?
M
M
10. Is the aircraft designed so that all landing
requirements are met (touch-and-goes,
crashes, etc.)?
M
M
Table 7 - Constraints Criterion Checklist
7. Objectives List
Below is a list of design objectives which MUNA Enterprises believes will closely approximate the
physical and flight characteristics of the aircraft:
OBJECTIVE DESCRIPTION
UNITS (IF
APPLICABLE)
VALUE
Aircraft Stall Speed
m/s (ft/s)
9.0 (29.5)
Flight Speed
Cruising Altitude
m/s (ft/s)
m (ft)
15.0 (49.2)
9.144 (30)
Maximum Estimated Payload
Empty Weight
kg (lb)
kg (lb)
2.73 (6.0)
0.939 (2.07)
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ENGI 7926: Detailed Design Report
Loaded Weight
kg (lb)
3.66 (8.07)
Target Payload Fraction
---
0.743
Wingspan
m (ft)
1.21 (3.95)
Propeller Size
mm (in)
304.8 x 152.4 (12 x
6)
Propeller Amp Draw
amps
22.0
Loaded Center of Gravity (from Shaft
Tip)
mmx,mmy,mmz (inx, iny,
inz)
175.21, 13.65, 0.11
(6.90, 0.54, 0.0043)
Tip-to-Tail Length
mm (in)
843.66 (33.22)
Maximum Theoretical Assembly Lift
(at 15m/s flight speed)
N (lbf)
35 (7.868)
Maximum Theoretical Assembly Drag
(at 15m/s flight speed)
N (lbf)
4.05 (0.91)
Table 8 - Objectives List
As seen in Table 8, the payload fraction target of .743 is very aggressive. MUNA Enterprises aims to use
an axisymmetrical payload which can be incremented when necessary in order to ensure maximum
flight stability. The fuselage can currently support a payload of 2 in x 2 in x 5 in volume, however the
payload used for flight will have smaller dimensions (approx. 2 in x 2 in x 2 in) to allow functional use of
the fuselage and variable shift of CGx.
8. Justification of Design Changes
The aircraft remains very similar in all respects to the concept which was created in the Research and
Concept Selection Report. The only distinct differences in the design would be the variation in fuselage
material, and the fuselage shape. ABS-M30 was chosen due to the demanding machining requirements
of the nose assembly for the fuselage. Using ABS-M30 and MUN's FDM machine, designing and
fabricating the nose of the aircraft was much easier than it would be if balsa was used as the material.
Constructing the nose assembly out of ABS-M30 also allowed for superior specific strength; in the event
of a crash landing, the nose can withstand a significant portion of the impact, limiting deformation of
structurally weaker components. Both sides of the fuselage are trussed framework configuration, and
the top and bottom fuselage sections are constructed from solid balsa sheet. The fuselage design
changed slightly when compared to the concept as well. Using the laser cutter for balsa framework,
further weight reduction and aerodynamic streamlining was made possible. The fuselage tapers down
toward its back end to an opening where the tailplane boom mount can be placed. This tapering feature
was not included in the concept, however it is a beneficial design characteristic due to its drag, material
and weight reduction.
9. Project Management Plan
With respect to the additional time allotted for presentation and report preparation, the project is on
task. When compared to the original Project Management Plan outlined in May 2012, the "Assembly"
17
ENGI 7926: Detailed Design Report
portion of Phase 2 is approximately 10 days behind schedule. Due to the additional requirements of
other courses undertaken by students this semester, Dr. Yang shifted the due dates of Presentation #2
and the Detailed Design Report, allowing teams extra time to complete design and analysis for their
aircrafts. The updated Project Management Plan shows plane assembly commencing as of Friday, June
29, 2012 and ending on Friday, July 6, 2012. Phase 3 (Testing) will commence on Saturday, July 7, and
end on Sunday, July 15. This revised timeline will still meet all milestones outlined by Dr. Yang. A table
of the visual indicators used within the Project Management Plan is available below. For additional
details, please see Appendix H - Project Management Plan.
Figure 12 - Project Management Plan Legend
10.
Conclusion
As a result of stringent analysis and thorough research, MUNA Enterprises' RC aero "Balsa Salsa" is now
fully designed. The "Balsa Salsa" is a high-wing configuration aircraft mainly constructed from balsa
wood, with a distinctive streamlined fuselage and a light empty weight of merely 2.07 pounds. In
combination with the attached bill of materials and technical drawing package, an RC hobbyist with
access to a machine shop can build an easy-to-assemble and lightweight aircraft well-suited for
recreational flying in a short period of time. The design is completely aligned with all current constraints
and specifications outlined by SAE. In its current state, the "Balsa Salsa" can compete in SAE Micro Class
competitions.
The innovation in the "Balsa Salsa" lies in its use of interlocking lightweight materials for rapid assembly
and unique combination of materials to achieve a lightweight aircraft with reputable rigidity and
optimum operability. The ABS-M30 nose assembly directly supplements the requirements of CG
placement and its high compressive strength assures the motor is kept intact - the most expensive part
of the plane.
MUNA Enterprises expects to produce its fully operable aircraft on July 7th, 2012. For more details,
please visit http://www.tinyurl.com/munaent.
18
ENGI 7926: Detailed Design Report
11.
References
1. Michael James. About.Com. “RC Airplane Parts and Controls”. About.com.
http://rcvehicles.about.com/od/rcairplanes/ss/RCAirplaneBasic_4.htm
2. RC Airplane World. “Understanding RC propeller size”. R/C Airplane World. http://www.rcairplaneworld.com/propellersize.html
3. Dr. James Yang. “ENGI 7926 (2012Spring) structure”. Memorial University Faculty of Engineering
and Applied Science. http://www.engr.mun.ca/~jyang/teaching/79262012/structure.htm
4. Stevens Institute of Technology Mechanical Engineering Department. “Micro Air Vehicle”.
Stevens Institute of Technology Mechanical Engineering Department.
http://www.stevens.edu/ses/me/fileadmin/me/senior_design/2006/group05/design.html
5. MUNA Enterprises. "Gantt Chart, Updated Thursday, May 24". Sawler Media.
http://www.sawlermedia.com/muna/wpcontent/uploads/2012/05/Gantt-Chart-UpdatedThursday-May-241.pdf
6. User “NFCR” (May 2012). “SAE RC Aero Micro Class Competition”. RCGroups.com.
http://www.rcgroups.com/forums/showthread.php?t=1649533
7.
Society of Automotive Engineers (2012). “SAE Collegiate Design Series”. Society of Automotive
Engineers. http://students.sae.org/competitions/aerodesign/about.htm
8. Dr. Leland M. Nicolai. “Estimating R/C Model Aerodynamics and Performance”. Lockheed Martin
Aeronautical Company, June 2009. Print.
9. Sadraey, Mohammad H. Aircraft Design: A Systems Engineering Approach. New York: Wiley,
2011. Print.
10. Hookey, Neil. ENGI 5932: Mechanical Vibrations - Class Notes. Canada: Memorial University of
Newfoundland, 2011. http://www.engr.mun.ca/~neil/6933/notes/notes.pdf
11. Voloch, Eduardo. Email. May – July 2012.
12.
Acknowledgements
MUNA Enterprises would like to thank Dr. James Yang for his assistance in preparation of this report,
along with Dr. Michael Hinchey for his assistance with stability analysis. Additional appreciation goes to
Mr. Eduardo Voloch for his help with stability analysis and overall design concepts.
19
ENGI 7926: Detailed Design Report
Appendix A - Foil Shapes
ENGI 7926: Detailed Design Report
Figure 13 - Selig 1223 and Eppler 420 Airfoils
Figure 14 - UI-1720 and CH-10 Airfoils
Appendix A - 1
ENGI 7926: Detailed Design Report
Appendix B - Balsa Tensile Stress Testing
ENGI 7926: Detailed Design Report
Material
Fracture
Figure 15 - Stress vs. Load of Balsa Wood
Figure 16 - Balsa Testing Apparatus
Appendix B - 1
ENGI 7926: Detailed Design Report
Figure 17 - Balsa Testing Apparatus
Figure 18 - Balsa Testing Apparatus
Appendix B - 2
ENGI 7926: Detailed Design Report
Figure 19 - Balsa Fracture
Appendix B - 3
ENGI 7926: Detailed Design Report
Appendix C - SAE Carrying Case Specifications Check
ENGI 7926: Detailed Design Report
The following image shows that all the components for the RC airplane easily fit within the specified box
dimensions.
Figure 20 - SAE Carrying Case Sizing Check
Appendix C - 1
ENGI 7926: Detailed Design Report
Appendix D - Drop Test Criteria
ENGI 7926: Detailed Design Report
Initial Drop Orientation Point of
Contact
Parameter
Value
Drop Height
~11m1
Gravity
-9.81 m/s
Nose of plane
Note:
1) Value is obtained from flight height multiplied by a safety factor. Drop Height = 30ft x 1.2 = ~11m
Table 9 - Nose Drop Test Details
Initial Drop Orientation Point of
Contact
Parameter
Value
Drop Height
~11m1
Gravity
-9.81 m/s
Bottom of plane
Note:
1) Value is obtained from flight height multiplied by a safety factor. Drop Height = 30ft x 1.2 = ~11m
Table 10 - Bottom Drop Test Details
Appendix D - 1
ENGI 7926: Detailed Design Report
Appendix E - Control Surface Servo Torque
ENGI 7926: Detailed Design Report
The following table shows the required servo torque required to move each control surface. It shows the
servo torques required for various angles of movement and various airspeeds.
Servo Torque Required for Various Intermediate Surface Deflections
Airspeed
9
13
18
(mi/hr):
Surface Servo
Angle
Angle Required Servo Torque (oz-in)
22
26
31
35
Aileron(s)
45
38
32
25
18
11
5
60
49
40
31
22
14
6
0
0
0
0
0
0
0
1
1
1
0
0
0
0
1
1
1
1
1
0
0
2
2
2
1
1
1
0
2
2
2
2
1
1
0
3
3
3
3
2
1
1
4
4
4
3
3
2
1
Elevator(s)
45
38
32
25
18
11
5
45
38
32
25
18
11
5
60
49
40
31
22
14
6
60
49
40
31
22
14
6
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
0
1
1
1
1
0
0
0
0
0
0
0
0
0
0
1
1
1
1
1
0
0
0
0
0
0
0
0
0
1
1
1
1
1
0
0
1
1
1
0
0
0
0
Rudder
Table 11 - Servo Torque Requirements
Appendix E - 1
ENGI 7926: Detailed Design Report
Appendix F - Engine Cover Heat Transfer Study
ENGI 7926: Detailed Design Report
Parameter
Flow Speed
Convection Coeff. For Air
Heat Source Temperature
Units
m/s
W/m^2
Kelvin
Value
15
5
383.15
Table 12 - Heat Transfer Testing Parameters
Figure 21 - Heat Transfer Testing Results
Figure 22 - Engine Cover Material Specifications
Appendix F - 1
ENGI 7926: Detailed Design Report
Figure 23 - Isoclipping from Heat Transfer Test
Isosurface of ABS-M30 region exhibiting temperature greater than 990 C (i.e. above Vicat Softening
Temperature)
Appendix F - 2
ENGI 7926: Detailed Design Report
Figure 24 - Heat Testing Results
Appendix F - 3
ENGI 7926: Detailed Design Report
Appendix G - Structural Analysis
ENGI 7926: Detailed Design Report
Figure 25 - Wing Structural Analysis
Figure 26 - Displacement and Von Mises Stress on Horizontal Stabilizer
Appendix G - 1
ENGI 7926: Detailed Design Report
Figure 27 - Displacement and Von Mises Stress on Vertical Stabilizer
Figure 28 - Landing Gear Stress Analysis, 52.0375 N Force
Appendix G - 2
ENGI 7926: Detailed Design Report
Figure 29 - Landing Gear Displacement Analysis, 52.0375 N Force
Appendix G - 3
ENGI 7926: Detailed Design Report
Appendix H - Project Management Plan