Project Report of Elective Studies
Ultralight Aircraft Design Philosophy
Submitted by: Khandoker Raisul AzadRegistration Number:
MAV-1310588003 Academic Supervisor: Eva Windbacher-Schwager Date of
Submission: 30/01/2014
PrefaceDifferent type of aircraft design has its own design
philosophy, such as the philosophy and process of commercial
aircraft as well as military aircraft are not same. The main
objective of this study is to represent the aircraft design method
in simpler way and find out the starting point where from a
designer can convulsion the journey without having previous
experience or expertise in aircraft design. It has been taken long
time to distinguish the starting point for the new designers,
because different aircraft design books and articles has taken the
approaches in different ways. As the number of aircraft
homebuilders are increasing rapidly as well as there are so many
passionate people I came across that they have the idea and
knowledge to design an ultralight aircraft but without having
proper guideline, they cant able to implement their innovative idea
and imagination. Consequently, it is being tried to find out the
simplest way of conventional aircraft design and this was working
as a prime intuition. So, after having done an extensive researcher
on this field, the whole development has got a skeleton, where the
designer need to follow the design ladders steps to provide
skeleton a workable shape.To solving this whole process who were
helping me, I do appreciate their feelings and knowledge towards
the aircraft design. I would like to acknowledge the authors, Dr.
Jan Roskom, Dr. Daniel P. Raymer, Chris Heintz, Johan D. Anderson.
Jr and Darrol Stinton who has written the incredible books for
aircraft design and after studying their books I have got the light
of aircraft design philosophy. I would like to express my gratitude
first to the almighty formerly the project supervisor Ms.
Windbacher-Schwager Eva. Then thanks goes to Mr. Hider Martin,
Late. DI Bruno Wiesler Head of Aeronautical Department, FH Joanneum
and Mr. Lukas Andracher for their incredible supports during the
project. They basically made me understood about aircraft design
and root out the confusions that I had so far.A Sincere thanks goes
to Mohammed Abdul Hamid, Head of Aeronautical Department, Prospects
College, UK, for his inspirations, suggestions and positive
feelings towards the project.
Ultralight Aircraft Design Philosophy (2014, Graz, Austria) Page
21
Contents1.Introduction62.Methods62.1.Identify the Key
factors62.2.Historical Data62.3.Literature Review72.4.Consultations
and Discussions72.5.Technological Resources72.5.1.Software73.Design
Philosophy84.Design Phase84.1.Requirements94.2.Initial
Drawing104.3.Weight Calculation104.3.1.Statistical
Rule104.3.2.Newton'sWeight Equation104.3.3.Per Square Foot
Method124.4.Engine and Propeller Selection124.4.1.Propeller
Selection124.5.Airfoil Selection134.6.Wing Design154.6.1.Tapper
Ratio: coefficient164.6.2.Wing Shape174.6.3.Dihedral
Angel184.7.Tail Geometry184.8.Fuselage Sizing204.9.Landing Gear
design204.10.Aircrafts Fragment Placement254.10.1.Engine
Location254.10.2.Tail Placement254.10.3.Wing Hinge
Point264.10.4.Landing Gear Location284.11.Control Surface
Design294.12.Further Weight Calculation304.13.Performance
Analysis324.14.Stability Analysis344.15.Detail
Drawing344.16.Prototype355.Results & Evaluation356.Conclusion
and Outlook:377.Summary388.References388.1.Books388.2.Reports,
Article, Thesis and Individual Papers:398.3.Electronic
Publications:398.4.Software399.Appendix A - Air properties and
conversions4010.Appendix B - Unit Conversions for
Design4111.Appendix D - Aircraft Materials Density4212.Appendix E
Relevant websites address42
List of FiguresFigure 1: Ultralight Aircraft Design Phase9Figure
4: Airfoil Selection flowchart14Figure 5: Wing Design Process
Flowchart15Figure 6: Wing Tapper Ratio16Figure 7: Mean Aerodynamic
Chord17Figure 8: Landing Gear Design Flowchart21Figure 9: Landing
Gear Primary Parameters22Figure 10: Wheel Load Geometry23Figure 11:
Landing Gear CG Location23Figure 12: Top and Side view of
Aileron29Figure 3: Wing structural Analysis32Figure 13: S_flapped
area33Figure 14: Aircraft CG Calculation28Figure 15: Detail Drawing
of Aircraft35
List of TablesTable 1: Materials weight in different part of
Aircraft12Table 3: Tail geometry Ratios19Table 4: Different
Aircraft Aileron data30Table 2: Empty weight fractions of different
materials31Table 5: C_fe Data33Table 6: Drag in different part of
Aircraft33Table 7: Lift enhanced data for flaps34Table 8: Design
parameters of different Aircraft35Table 9:Design parameters of
certain aircraft37Table 10: Aircraft performance data37
IntroductionAircraft design has its own philosophy which is
being differ from designer to designer, with their knowledge,
experiences and expertise. It is very difficult to come up with a
feasible aircraft design if the designers philosophical view is not
rich enough. The philosophical view has its own and distinctive
influences during the implementing level to sophisticated level of
aircraft design. To improve and improvise the aircraft design
procedure as well as find out the acceptable and easy to understand
for all groups of aircraft designers, this study has taken place.
The methods has been described in this study can mostly be used for
the conventional aircraft design. Rather than that, in terms of
conceptual approach the whole information of this study can be used
to develop the idea of particular designers. Perhaps, it is very
difficult to come up in a unified design process for aircraft
design because the designers have to change his approaches time to
time for the situation demand. So, every designer has his/her own
idiosyncratic design approaches but here, its been tried to come up
with a new method from the blend of all existing approaches to
embellish the aircraft design process. The project that is being
undertaken was not just an academic objective but also a choice at
a personal level. Thus the motivation lied in the just for the fun
of it factor.
MethodsDifferent methods had been followed to find out the
correct philosophical view for ultra-light aircraft design which is
being described as follows. Identify the Key factorsThere are many
different methodologies which can lead to satisfactory design
solution. But among them which would be paramount for particular
project, is the main objective of this part is to sort that out.
So, it is indeed be crucial to find out the key factors, which may
play constructive or adverse role in the whole design session, then
one can brush off the unnecessary thing from the project. Whereas
the optimized output would be acquired to achieve the expected
uttermost peak.Historical DataAircraft design output is very
vulnerable, like, from a good guesstimates and calculations of the
different phase of Aircraft design, may produce some unexpected
output value which would not be matched up with the calculated
value. Perhaps, for an inexperienced designer it would indeed be
best to try something usual and quite similar with successful
previous design. To testify the results, that the applied methods
are correct, some successful design parameters of different
aircraft of past has been taken and it is being observed that most
of the cases the theory and equations are successfully testified
with minimum tolerance. E.g. Thorp T-18 C (Sunderland Aircraft),
Crawdad (Foot launched motor glider).Literature ReviewIn this
journey, literature review helped to grasp the basic concept of
aircraft design. Furthermore it helps to dig down more about the
flight physics and other factors effecting the design philosophy.
In addition, it is being schooled about the proper use of Articles
and scientific papers. As there are so many different approaches
are for aircraft design, and literature review helps to know detail
of those approach and understand them as well as apply them where
they are being needed. Like Dr. Jan Roskom has written about 8
volume of Aircraft design book but on the other hand Dr. Dan Raymer
writes Simplified Aircraft design for Homebuilder which covers
mostly all the requirements of aircraft design. But, for
implementation it is being indispensable to follow different books,
articles and scientific papers.Consultations and DiscussionsSome
parts of this project is really difficult to understand without the
help of an experience person. When something comes like that,
instantaneously contact has been made with the project supervisor,
the professors and the fellow mates who worked with aircraft design
before. Perhaps it is been wonderful to consult with those
personnel for clarifying the idea.Technological ResourcesTo find
out the project quarries, technological resources is needed to make
it more simplified and time constrain. The main objective was
behind this approach to use available web based resource which is
free of cost, but it took bit time to get use to implement them
successfully in the research work. There are more or less all
information can be found in the web but the designers need to know
where and to implement them in effective way.1.1.1. SoftwareThere
are some software is still available in Internet which is really
helpful to get the output of the working project within short
period of time, which may help the designer that he is walking in
the right direction or not. For this project there are different
software is being use: XFLR5 Martin Hollman aircraft Software.
Raymer SoftwareXFLR5 is one of the software who can save the
designers time. By using this software designers can know the
following parameters which is being needed for the calculation of
whole design work. ClMax Lift vs Drag Curve Stall Speed Stability
and control analysis 2D and 3D analysis Polar object Airfoil
modification Inertia estimations Viscos and incised calculation
Neutral point Centre of pressure Static margin, etc.
Design PhilosophyFor designing an aircraft the designer has to
estimate everything, nothing is concrete here. Moreover, a designer
can perform a good design, if he has the adequate knowledge and
expertise. The best aircraft design is Keep it short and simple
(KISS). Aircraft controlling systems and processes need to be
designed in such a way that they are as simple as possible would
parallel with the task required for them. The design must mirror
the both goodergonomics and a conscious effort to minimize human
factors. Encase of Ultralight Aircraft Design air safety should be
taken into in account by introducing precise and simplified design
methodology to produce real time response in terms of abnormal
behavior of aircraft. It should be kept in mind of a designer that
the controlling system supposed to be simple and easy to adoptable
for the pilot. Perhaps it is an intellectual act and for performing
that knowledge, practice and experiences is being needed.
Nevertheless, to strengthen the philosophical view about aircraft
design every single thing should be taken into account. Design
PhaseThis is the important phase and the final production quality
as well as cost effectiveness is depend upon it. Conceivably, the
more precise and specific design phase would be the more tuned and
finished product would have been achieved. Which may help to save
lot of cost and labor. So, to produce a fine tuned out put this
phase is being divided into different wings. Figure 1: Ultralight
Aircraft Design PhaseN0Ultralight Aircraft
RequirementsIt is easy to get the likely requirement, if the
designer ask himself with what? And why? Perhaps these are the big
questions to set a specification for the aircraft. This process is
not only applicable for Ultra-light aircraft but also every kind of
aircraft design including commercial and fighter aircraft. But it
is indeed to think that would be made as a commercial product or as
a homebuilt product. However, Ultralight Aircraft is mostly
designed and manufactured as a homebuilt product. Requirements
setting is being performed as per the goal setting of the designer.
Requirements/Specification is working as the gate way to reach up
to the projected goal. It is the initial starting to get into the
deep design process. The following specification might be followed
to get a notion about the further processes- Range Max Cruising
Speed Min Stall Speed Total Weight Cruise Height Engine Lycoming
Wing Area Wing Chord Wing SpanInitial DrawingTo visualize the
design a sketch is necessary for following up the projected design.
It will help to change some initial decision as well as
optimization. As the whole study is about conventional aircraft
design, in this phase designer may have concentrate on minimum
requirements. Weight CalculationThis is the pilot part of the
design, the more augmented by the designer this part the more
successful design outcome may achieved. But it should be remembered
that the main objective of Weight Estimation is to get an appraised
value and depending on it, other factors can be calculated.1.1.2.
Statistical RuleBut the statistical rule would have been another
interesting and brilliant approach to do this, it is such that, the
designer should take some aircraft data and which is similar or
almost similar characteristics or specification and does have the
successful flying records. Often statistical weight relationship
follow an exponential equation. This means that if a graph is being
plotted there would be straight line depending on those values.
Therefore, depending on their statistical records the Weight
Calculation process can be performed. Furthermore, some other
factors can be taken into consideration. Perhaps the faster
aircraft have a higher weight per square meter than the slower
aircraft. A weight per square meter verses maximum speed graph
could be drawn and can be used for weight prediction of particular
aircraft. The higher power engine has a lower weight ratio. Most of
the cases the home builders do consider the weight of landing gear
is about 5% of empty weight. But it will depend on particular
requirements that where the aircraft would land. The weight
estimation involves a component build-up, in much the same fashion
as we measured aircraft drag.1.1.3. Newton'sWeight Equation The
following rule can be followed for the weight calculation:Using
Newton'sweight equation:w = m g {m= Mass & g= Gravitational
force} To find out the m we need to know the density. The mass of a
discrete component can be calculated if we know the size of the
component and its material composition. Every material (iron,
plastic, aluminum, gasoline, etc.) has a uniquedensity. Densitydis
defined to be the mass divided by the volumev: So,d = m/v { m= Mass
& v = Volume}or, m = r v To find out the Volume-v the following
formula can be followed.
Cubeside3
Rectangular Prismside1 side2 side3
Sphere(4/3) pi radius3
Ellipsoid(4/3) pi radius1 radius2 radius3
Cylinderpi radius2 height
Cone(1/3) pi radius2 height
Pyramid(1/3) (base area) height
Torus(1/4) pi2 (r1 + r2) (r1 - r2)2
Now, we can rewrite the Aircraft weight equation: w = m g v.As
aircraft is being made of different component. So the Total Weight
would be:W = w(fuselage) + w(wing) + w(engines) + w(payload) +
w(fuel) + ..For Ultra-Light Aircraft Design the following
components of aircraft should be considered to get the total
weight: Wing Horizontal Tail Vertical Tail Fuselage Landing Gear
Surface Controls Propulsion System Instruments and Navigation
Electrical System Electronics Crew Payload Fuel1.1.4. Per Square
Foot MethodThere is another way to estimate the weight of aircraft
and it is called the Pound per square foot method. Conceivably this
method may give a better result than the fancy equations. The
following table is based on various ultralight and homebuilt
aircraft.Weight EstimationWingHorizontal TailVertical
TailFuselage
lb/sq-ftlb/sq-ftlb/sq-ftlb/sq-ft
Metal1.1 to 2.00.9 to 2.00.9 to 2.01.2 to 1.4
Fiberglass1.6 to 2.20.9 to 2.00.9 to 2.01.2 to 1.4
Carbon fiber1.2 to 2.00.9 to 2.00.9 to 2.00.7 to 1.2
Fabric1.0 to 2.00.8 to 1.50.8 to 1.51.4 to 1.8
Table 1: Materials weight in different part of AircraftEngine
and Propeller SelectionIf the weight is being estimated properly
the Engine Selection process would be easier. Now this is the turn
for calculating the power needed to run the estimated weight and
make it fly by working against the gravitational force. To do so
engine itself cant do anything without a propeller, so its
necessary to select suitable propeller, which can perform
accordingly with selected engine. Following formula may help to
selecting engine and propeller.Ultralight: W/hp = 325 Vmax-0.75By
using this formula required power loading can be estimated on
desired maximum speed. Power loading typical range from 10 to 15
per Hp for homebuilt aircraft.Horsepower Required: Hp= W0/ Power
Loading {W0= Empty Weight}1.1.5. Propeller SelectionThe designer
need to know the propeller diameter in order to know the thrust
producing by the particular propeller.Propeller Diameter: D=
22(hp)1/4 {diameter in inches} (for 2 bladed Prop)Propeller
Diameter: D= D= 18(hp)1/4 {diameter in inches} (for 3 bladed
Prop)It is needed to be considered during propeller selection, if
the propeller is too large the tips may approach sonic speeds in
high speed flight causing a loss of thrust and huge amount of noise
would have been produced. To avoid this the tip speed may calculate
by the following formula:Tip Speed: Vtip= {D in ft and V in
ft/sec}Airfoil Selection Wing is the lock of an aircraft and
airfoil is the key of the lock and without having this two things,
it is impossible to get into the ecstasy of aircraft design.
Perhaps, it is indeed very important to select the suitable
airfoil, which do has effect on aircraft lift, drag, weight and
stability near the stall. In general a airfoil will generate more
lift if it is thicker, has a more round edge and has more camber.
Airfoil with more camber tend to have a greater pitching moment and
for compensating this, more trimming force is required. With a
sharp leading edge airfoil may have less drag and it is more prone
to have sudden stall.Some airfoil is design such that encourage the
laminar flow and front to rear air flows without having disturb. In
terms of structure and weight a thick airfoil is not suitable,
there for airfoil with a lot of camber is good for lower speed but
have extra drag in high speed flight. On the other hand uncambered
airfoil is good at high speed but does produce lots of drag in
terms of low speed flight. Airfoil can be designed as per the
requirement but it would be bit difficult and time consuming, other
than that a designer can chose from the wide collection of airfoil
from some website, like
http://aerospace.illinois.edu/m-selig/ads/coord_database.html,
those data from this website can be used in different software and
modified as per the requirements.
Figure 2: Airfoil Selection flowchartThere is another thing need
to consider during the airfoil selection, airfoil supposed to place
at some incidence angel to the fuselage so that they are at the
correct angel of attack for creating the lift, which is being
needed during cruise, with the fuselage at zero angel of attack so
that it doesnt create unwanted drag. Calculate the lift coefficient
during cruise by following the below equation, then find the angle
of attack () that lift coefficient.Wing Loading: W/S=qCL, Where, q=
Vmax2 And, CL= CLmaxSo, cruise lift coefficient: CL-cruise= Angel
of attack: = CL-Cruise{10+18/A}+zeroliftOr, If the wing is Swift: =
CL-Cruise + zeroliftHere, A= Wing Aspect Ratio. zerolift will come
from the selected airfoil data. An airfoil with camber makes lift
even at zero angel of attack (AOA), so the calculated angel of
attack is too large. In the airfoil lift vs. angel of attack, find
the AOA thats gives Zero lift and this supposed to be a negative
number. If an airfoil used with a lot of camber, there is nothing
to be surprised if the adjacent incident angel is very small.It is
very difficult to get the value of CL and CL-Max and the manual
calculation is very difficult and time consuming, perhaps that can
be attained by using software otherwise choosing NACA or other
airfoil, which CL-Max value is available. Because it is very
significant for the 2D and 3D lift calculation.Now question arises,
what would be the twist angel? Which is related to wing, mostly the
wing twist for the ultra-light and homebuilt aircraft is 2-3
degrees, with higher airfoil incidence at the root and a lower
incident at the tip. Which is known as washout and makes the wing
stall first at the root because it has a higher AOA.The incident
angel calculated above is the airfoil angel at the Mean Aerodynamic
Chord (MAC), not the centerline wing root airfoil or the airfoil at
the side of the fuselage. The above things can be done by XFLR5
software with an acceptable accuracy and short period of time.Wing
Design
Wing design is quite challenging part of aircraft design, this
part and airfoil selection parts importance is vice versa. For
designing a wing, lift calculation is the key part because it is
need to know that the produced lift by the wing can be compensated
by the aircraft weight or not. And for lift calculation CL-max need
know is very important but it is very difficult to calculate the
CL-max from the airfoil (2D) to the wing (3D). To do so, XFLR5
software can be chosen. So wing design process can be followed by
the bellow flowchart.Figure 3: Wing Design Process FlowchartFor
calculating take off gross weight, wing area should be calculated
by the following formula.Wing Area: S = W0/W/S { W0= Take off gross
weight }Wing span is the main factor in drag due to lift. A larger
span results in lower drag due to lift. But a larger span is
comparably heavier, so picking up a suitable value of wing span,
all factors should be compromised.CL-max Calculation:CL= CL-max
this is been discussed before as an important factor which is bit
complex to define the value in terms of calculation. But there is
another easiest way to find the particular value, if weight is
considered equal to lift W = L. From the lift formula decision
Lift: L = V2S CL Or, W = V2S CLOr, CL = V2S WFor a typical
ultralight aircraft, CL-max will be about 1.4 without using flaps.
1.1.6. Tapper Ratio: coefficient Tapper Ratio () is another
important geometric parameters in terms of wing design. Tapper
Ratio = Root Chord/ Tip Chord. Tapering of a wing is used mostly to
change the span wise lift distribution. A wing with no tapper has
too much lift out near the tips. Wing tapper may also reduce the
structural weight, because the root chord is longer than the tip
chord, which provides a greater leverage for handling the bending
moments. This may allow the spars and skins to be thinner. The
tapper ratio is about 0.4 will usually provide the best compromise
between aerodynamics and structural weight, and this can be apply
for larger aircraft with upswept wing. For ultra-light aircraft
this weight savings due to tapper may not apply since the skins can
only be so thin. Also the ultra-light aircraft may experience the
tip stall if the tip chord is so short. It is therefore suggest
that the tapper ratio suppose not to be less than 0.5 (tip chord
equals half the root chord).
Figure 4: Wing Tapper Ratio1.1.7. Wing ShapeNow, by using the
following formula the wing shape can be visualized.Wing Span: b =
Root Chord: Croot = 2S/b (1+) { = Tapper ratio}Tip Chord: Ctip =
CrootNow the next step to be considered about make the aircraft
stable. That is why the aircraft is design in a way the center of
gravity and the wing are in the right location in respect to each
other. And to do so, Mean Aerodynamic Chord MAC is needed to find
out and this will be needed for the CG calculation as well. At the
root of the wing, draw a line parallel to the centerline of the
fuselage and extending forward from the leading edge as well as
rearward from the trailing edge. Both lines should be the same
length of the tip chord. The same thing supposed to be done in
terms of the tip but drawing the lines the length of the root
chord. Now, connect the ends of the lines so that they create an
"X" over the wing panel. Where the two lines intersect is the span
wise location of the MAC.
If the plan indicates that the CG should be located at some
percentage of the MAC, then measure the MAC and put the CG the
given percentage back from the leading edge along the MAC. As a
instance, if the MAC is 10" and the plan indicates the CG should be
25% back from the leading edge, then the CG is 2-1/2" back from the
leading edge at the MAC.Figure 5: Mean Aerodynamic ChordThe lines
cross at the span wise location of the MAC. It is not the fore/aft
CG location (unless the CG happens to be located at 50% MAC).The
following formula will give the measurement of MAC. = Taper Ratio =
(Tip ChordRoot Chord)MAC: = Croot x 2/3 x ((1 + + 2)( 1 + ))1.1.8.
Dihedral AngelThe principal aim for applying a wing dihedral is to
enhance the lateral stability of the aircraft. The lateral
stability is primarily the tendency of an aircraft to return to its
original trim level-wing flight condition if disturbed by a gust
and rolls around the x-axis.
Figure 6: Effect of dihedral angel before and after gust.When
the dihedral angle is applied on a wing, the wing effective plan
form area (Seff) is reduced. This basically reduce the aircraft
total lift, which is undesirable. So, it is being suggested that to
consider the lowest value for the dihedral to minimize the lift
reduction. The effective wing plan form area as a function of
dihedral angle is determined as follows.Effective Wing Plan Form
Area: {Here, Note: Typically dihedral angle is being selected
between15 to +10 deg.Tail GeometryAircrafts have tails to make
moments, moments are made by having some force act at a distance
around the point of rotation. For tails, the moment arm is measured
from the MAC of the wing. It should be remembered that the tails
are used for balancing the aircraft instead of producing lift, in a
sense it produce lift but in downward. By using the following
equations the horizontal and vertical tail size (Surface Area) can
estimate.Horizontal Tail: = CHT (/LHTVertical Tail: = CVT
(/LVTHere,LHT = Tail moment arm length (for Horizontal tail)LVT =
Tail moment arm length (for Vertical tail)b wing = Wing Spanwing =
Wing MAC
CHT and CVT are coefficient of this equation. Typical ultralight
aircraft values can be used o.5 for horizontal tail and 0.04 for
the vertical tail.The aerodynamic fore generated on the tail
readily change direction whether the aircraft yawing right or left
and/or pitching up or down and depending on which direction rudder
or elevator are deflected. So, it is unusual to use camber airfoil
rather symmetric for the tail, mostly NACA 0012 airfoil is being
used for the tail section (Horizontal and Vertical). As it
concerned that the higher aspect ratio of the wing is more
aerodynamically more efficient than the lower aspect ratio. So, it
is better to take lower aspect ratio than wing while tail design.
Because when the wing stall the already have some control
authority. The following formula can be used.Horizontal Stabilizer
Span: {Here,}Horizontal Stabilizer Root Chord: {Here, }Horizontal
Stabilizer Tip Chord: Now, time to calculate the Vertical
Stabilizer staff. Likewise the (Aspect Ratio of Horizontal
Stabilizer) the value of should be imagine. Mostly 1.3 to 2.0 is
being choose by designer. Perhaps the following table (3) can also
be followed.Vertical Stabilizer Span: {Here,}Vertical Stabilizer
Root Chord: {Here, }Vertical Stabilizer Tip Chord: For getting the
full idea about the tail aspect ratio, taper ratio, and sweep need
to select. They are not as critical as the wing, and its OK just to
make the tails so they look like tails, provided they have the
right area. Following table can be help to develop the notion.Tail
GeometryHorizontalVertical
Aspect RatioTaper RatioAspect RatioTaper Ratio
Conventional3 to 50.3 to 0.61.3 to 2.00.3 to 0.6
T-tail3 to 50.3 to 0.60.7 to 1.20.6 to 1.0
Sailplane6 to 100.3 to 0.51.5 to 2.00.4 to 0.6
Table 2: Tail geometry RatiosElevators for ultralight aircraft
are usually about 45% of the tail chord. Rudders are normally about
40% of the tail chord. The vertical tail plays a key role in spin
recovery and the horizontal tail can hurt its
effectiveness.Fuselage SizingHow big the fuselage would be? The
answer of the question depends upon the designers initial
requirement, like the total people, payload and the range of
aircraft. The designer should consider the comfort, on the same
time provide an eye catching shape which will produce less drag. By
using the following equation a rough approximation about the length
of fuselage.Fuselage Length: L = 3.6 0.23Of course this estimation
should only be considered as a starting point. There is
considerable debate about the best value for fuselage finesse ratio
(Length/diameter). Numerous design books such as the classical
Hoerner Fluid Dynamic Drag say that the lowest drag occurs when the
finesse ratio is around 3. However, most airplanes have much higher
finesse ratio. But it is true that for ultralight aircraft finesse
ratio 3 is most suitable.Landing Gear designLanding gear is an
important factor for landing rather take off and which is named
considering the use and purpose. So, landing gear is needed to
design such a way that the maximum amount of energy is being
absorbed by this without affecting the aircraft structure.
Basically the landing gear design mainly depends on the weight and
stall speed of the aircraft. So, Landing gear should be able to
bear 90% of the weight of the aircraft while standing. Mostly the
landing gear of an ultralight aircraft is fixed and tricycle.
Perhaps the following parameters should be considered in terms of
landing gear design. Type (e.g. nose gear (tricycle), tail gear,
bicycle) Fixed Height Wheel base Wheel track The distance between
main gear and aircraft cg Strut diameter Tire sizing (diameter,
width) Load on each strut
Figure 7: Landing Gear Design Flowchart
Figure 8: Landing Gear Primary Parameters
In terms of design procedure, the landing gear is the last
aircraft major component which is designed. In another word, all
major components (such as wing, tail, fuselage, and propulsion
system) must be designed prior to the design of landing gear.
Furthermore, the aircraft most aft center of gravity (cg) and the
most forward cg must be known for landing gear design. In some
instances, the landing gear design may drive the aircraft designer
to change the aircraft configuration to satisfy landing gear design
requirements. The primary functions of a landing gear are as
follows: To keep the aircraft stable on the ground and during
loading, unloading, and taxi. To allow the aircraft to freely move
and maneuver during taxing. To provide a safe distance between
other aircraft components such as wing and fuselage while the
aircraft is on the ground position to prevent any damage by the
ground contact. To absorb the landing shocks during landing
operation. To facilitate take-off by allowing aircraft acceleration
and rotation with the lowest friction.In order to allow for a
landing gear to function effectively, the following design
requirements are established: Ground clearance requirement Steering
requirement Take-off rotation requirement Tip back prevention
requirement Overturn prevention requirement Touch-down requirement
Landing requirement Static and dynamic load requirement Aircraft
structural integrity Ground lateral stability Low cost Low weight
Maintainability ManufacturabilityWheel base (B) plays an important
role on the load distribution between primary (i.e. main) gear and
secondary (e.g. nose, or tail) gear. This parameter also influences
the ground controllability and ground stability. Thus, the wheel
base must be carefully determined and an optimum value needs to be
calculated to ensure it meets all relevant design requirements.
Figure 9: Wheel Load Geometry
Calculation of the static loads on each gear is performed by
employing equilibrium equations. Since the aircraft is in static
equilibrium, the summation of all forces in z direction must be
zero:Furthermore, the summation of all moments about o is zero:
Thus the percentage of the static load (i.e. aircraft weight)
which is carried by the nose gear is:
In addition, the percentage of the static load which is carried
by the main gear is:
In the case of a tricycle landing gear, the load on the main
gear is divided between left and right gear, so each wheel will
carry one half of the main gear load (i.e. Fm).The above-mentioned
relationships are applicable only in static situations. There are
two other interesting conditions that cause landing gear to
experience different loadings: 1. Change in the aircraft center of
gravity location; 2. Dynamic loading. Due to the possibility of a
change in the load distribution, or having different combinations
of cargo, or number of passengers, the gears must carry a load
other than the nominal static load. In the x-axis, an aircraft
center of gravity is allowed to move between two extreme limits: a.
most aft location (Xcgaft), and b. most forward location
(Xcgfor).
Figure 10: Landing Gear CG LocationFor tricycle configuration
with most aft and most forward cg locations. The following
equations govern the minimum and maximum static loads on each
gear:
Furthermore, landing gear tends to experience a dynamic loading
due to aircraft acceleration and deceleration during take-off and
landing. The nose gear will have to carry a dynamic loading during
the landing operation when aircraft is braking. During braking
segment of the landing operation, the following equilibrium
equation may be written:
Where aL is the braking deceleration and g is the gravitational
acceleration. Therefore the nose gear load is:
The first term of equation 9.13 is the static load, but the
second term is referred to as the dynamic loading:
Hence, the total load on the nose gear during landing will
be:
To insure the ground controllability in a tricycle landing gear
configuration, the parameter Bmmin should be greater than 5 percent
of wheel base and the parameter Bmmax should be less than 20
percent of the wheel base. These equations and requirements are
employed to determine wheel base plus the distance between cg and
nose gear, and cg and main gear. With a similar approach, the
dynamic loading on the main gear during take-off acceleration with
an acceleration of aT will be determined as follows:
Thus, the total load on the main gear is:
These static and dynamic loadings are utilized in determining
nose and main gears locations, strut load, and wheel and tire
design. It must be noted that the main gear is usually carrying a
total load which is greater than the aircraft weight.Aircrafts
Fragment PlacementAfter designing the different part of the
aircraft, the question arises that how and where should be placed
and install them. In this section of this research those things
will be identified.1.1.9. Engine LocationIn general, engine
location selection is an important part, because many factor is
related with that. From them most important factors are- Stability
Performance Aerodynamic interference Landing gear location Aircraft
CGConsidering the above factors the designer should decide where
should be the engine positioned. Because there no equation for
that, it depends upon the general understanding and requirements of
the designer.But considering the ultra-light aircraft design mostly
Tractor or Pusher type propulsion system is being selected. Both of
them have advantages and disadvantages. So, after selecting the
engine location it is needed to install them on the centerline of
the fuselage, which does have a great impact on aircraft
stability.1.1.10. Tail PlacementThe tail should be placed
sufficiently far back of the fuselage that at stall the wake of
horizontal tail does not mask the rudder on the vertical tail. In
terms of tail placement, it is need to be remembered that it
supposed to keep as much as far from the wing position to get more
balance as well as structural strength would have been considered.
As the size of the fuselage is already being defined, so it is
suggested to keep it the most aft part of the fuselage, though
conventional design is being considered. 1.1.11. Wing Hinge
PointNow question areas that where should be the wing would hinged
to get optimum balance during flight and ground maneuvering? Hence,
to find the specific position Center of Gravity (CG) should be
calculated differently for the fuselage and wing. To do so some
other factors would be calculated and that is shown below.
Figure 11: Aircraft body and wing CG locationEstimated CG:
Here,
So, from the above equation the estimated location can be found
for the wing where it can be placed. But to optimize the hinge
position the aerodynamic center should have been considered which
lie behind the aircraft CG. Aerodynamic Center of the wing
body:
Static Margin: SM = Here, {This is being measured in wing
Design} CG location of the aircraft
The SM is a simple and direct measure of stability. 12% to 20%
(ie. 0.12 to 0.20) is expected for nice flying and stable design.
These suggested values include an allowance for a propeller in
front which is destabilizing, if pusher propeller is being used it
supposed to be reduced by 3% to 5%.So now the Static Margin
equation can be rewrite, Again: SM = Hence, from the aerodynamic
center of the wing body equation, it can be assumed that
Aerodynamic Center of the wing body (Wing Fuselage) = Aerodynamic
Center of the wing [wing]. Also for the simplicity it can be
considered that. So, considering the above situation the
Aerodynamic Center of the wing body can rewrite as
follows.Aerodynamic Center of the Wing: wing = However, the win
would be placed in such location that aerodynamic center is {wing =
[calculated value]} behind the nose of the aircraft. Note: Center
of Gravity CG can be calculated by following formula or there are
some website where the CG calculator is freely available
(http://adamone.rchomepage.com/cg_calc.htm). To get more idea about
CG calculation the following figure may help out.
Ref: http://www.grc.nasa.gov/WWW/k-12/airplane/acg.html
Figure 12: Different objects CG calculation1.1.12. Landing Gear
LocationAs the landing gear is already selected and designed, so
now this is the time to place it in the correct position to get
optimum control, while the takeoff, landing and ground maneuvering.
To identify the landing gears positions, it is indeed be important
that find out the whole aircraft CG, which already be calculated in
the part (4.10.3), because CG is the brainchild of aircraft
performance and attachment of different fragments of aircraft.
Hence, the following equation can be followed to ascertain the
location of landing gear. Main Landing Gear Position: )Here,
Coefficient of lift during takeoff.It is suggested that the main
landing gear should carry 85% of the a/c weight and the nose
landing gear 15%.Nose Landing Gear Position: Typically, for
standard value can be considered.
Figure 13: Landing gears positionControl Surface DesignAs the
ultralight aircraft is the concern, so the control surface would be
limited with Flap, Aileron, Elevator and Radar, within them aileron
design importance is greater, because the primary function of an
aileron is the lateral (i.e. roll) control of an aircraft; however,
it also affects the directional control. Due to this reason, the
aileron and the rudder are usually designed concurrently. The
deflection of any control surface including the aileron involves a
hinge moment. The hinge moments are the aerodynamic moments that
must be overcome to deflect the control surfaces.In terms of
aileron design, four parameters need to be determined. They are: 1.
Aileron plan form area (Sa); 2. Aileron Chord/Span (Ca/ba); 3.
Maximum up and down aileron deflection (+ or - Amax); and 4.
Location of inner edge of the aileron along the wing span (bai).As
a general guidance, the typical values for these parameters are as
follows: Sa/S = 0.05 to 0.1, ba/b = 0.2-0.3, Ca/C = 0.15-0.25,
bai/b = 0.6-0.8, and Amax= 30 degrees. Based on this statistics,
about 5 to 10 percent of the wing area is devoted to the aileron,
the aileron-to-wing-chord ratio is about 15 to 25 percent,
aileron-to-wing-span ratio is about 20-30 percent, and the inboard
aileron span is about 60 to 80 percent of the wing span. Table
12.17 illustrates the characteristics of aileron of several
aircraft.
Figure 14: Top and Side view of AileronTherefore the following
statistical data of different aircraft would develop some idea on
aileron designing.
Table 3: Different Aircraft Aileron dataFurther Weight
CalculationOtherwise to calculate the starting weight of an
aircraft that can just exactly make the range requirement, which is
called Takeoff Gross Weight W0. To calculate that, it should
estimate the Parasite Drag Coefficient CD0 and this the part of the
drag that does not change when the lift changes. Another drag K
drag due to lift factor it helps to estimate the drag on the wing
caused by the creation of lift. The drag due to lift coefficient K
time the square of the lift coefficient. CD0 is mostly related to
total wetted area or surface area- Swet of the particular design
including the top and bottom of the wings, the top, all sides and
bottom of the fuselage and both side of the tails.The wetted area
can be estimated using a ratio to the wing area Swet/Sref, here,
Sref- Reference wing area. Since the wing area is defined as the
top view projected area, the wetted area must be at least twice the
wing area. Actually, even for a pure flying wing the wetted area is
larger than two due to the area around the leading edge. The
following equation can help to find the Parasite Drag
Coefficient.Parasite Drag Coefficient: CD0 = Cfe (Swet/Sref ) {Cfe
= Skin friction Coefficient} Generally, for Conventional Design
Swet/Sref 3.8 for single engine and 4.6 for twin engine. Cfe for
average design- 0.0090 and for smooth design- 0.0065 and those
values are applicable for Single engine fixed landing gear. Now,
Drag due to lift factor: K = 1/0.75A = 0.424/A {A = Aspect
ratio}For most Ultralight aircraft the A is somewhere 6 to 8.
Perhaps a higher value gives lower drag and therefore more range
and climb rate but it is usually heavier and may reduce roll
response.Now, Lift to Drag Ratio: L/D = The fuel burn of the engine
is expected as engine specific fuel consumption. This typically o.4
to 0.6 pounds of fuel used per hour per horsepower. But the value
0.45 can be use for most modern aircraft piston engine. So, to
convert per pound of fuel per second per horsepower produced, use
0.45 divided by 3600, or a specific fuel consumption of Cbhp =
0.00013 lbs/sec/bhp.Now, find out fuel fraction, which is known as
Breguet equation to calculate the remaining weight of aircraft
after the cruise.Fuel Fraction: Wf/W0 = 1- 0.975
e-RCbhp/550pL/DHere, the 0.975 term is used an approximate
allowance for additional fuel used during takeoff, climb, descend
and landing. The value of e=2.7183approximately. R is the range in
feet.It is being needed to determine the empty weight, which is
estimated as a fraction of the takeoff weight (We/W0). We includes
everything other than fuel, people and payload. We fraction is a
non-dimensional ratio that doesnt change much for different
aircraft design.Empty weight fraction: We/W0 = a W0-0.09 {a=
Materials Type, which is given in following table}a Material
TypeSingle EngineTwin Engine
Metal or Wood Design1.191.40
Composite Design1.151.35
Table 4: Empty weight fractions of different materialsThe sizing
equation below calculates the aircraft weight W0 that just meets
the range requirement R. The weights of the people and payload come
from the designer requirements, but a normal weight allowance for
people is around 80 to 90 Kg.So, W0= Wpeople+ WPayload/1- We/W0-
Wf/W0It is indeed be important to calculate the thickness of the
structural parts required to safety withstand the expected loads,
including a factor of safety.
Figure 15: Wing structural AnalysisIn the above diagram the
total wing lift L equals the aircraft weight W time the load factor
n, (L=nW). It is being assumed that the lift is uniformly spread
across the wing, so the total lift acts on average, at appoint half
way out of span, this creates a moment at the wing root, which
equals the lift on one side (nW/2) times the moment arm distance.If
the lift moment equals the wing root reaction moment (T=C),
then,Equating Moment:, So, C = T =Performance AnalysisIt is the
part where the designer supposed to prove the estimation that has
been done so far. But, it is true that there would some tolerances
between the calculated value and the estimated value, perhaps if
the tolerance is so high, then it would indeed be redo the whole
thing would be suggested.Stall Speed: VStall = {Here, = air density
at sea level, W = Weight, S = Wing Area}Takeoff Parameter: T.O.P =
1.21{Here, HP = Horse Power, W = Weight, S = Wing Area}Rate of
Climb: Vv=V{(Here, T = Thrust produced = ) {For the above Thrust
Calculation the designer should calculate the thrust two times,
once at 100% power and once 75% (Cruise Power Setting).Total Drag:
DTotal = S(CD0 + KCL2) Here, CD0= Cfe [Cfe= Equivalent skin
friction coefficient, Sref= Wing reference area.CfeSingle
EngineTwin EngineSail Plane
Average Metal Design0.00580.00480.0038
Smooth Composite0.00500.00450.0030
Table 5: C_fe DataDrag AreaDrag/Dynamic Pressure (D/q) per unit
frontal area
Exposed Wheel & Tire0.25
2nd wheel in Tandem0.15
Streamlined wheel and tire0.18
Wheel & Tire pants0.13
Round Strut0.30
Streamlined Strut0.05
Flat Spring Gear Leg1.40
Fork or Irregular Fitting1.0 to 1.4
Speed brake- Fuselage1.00
Speed brake- Wing1.60
Windshield- Smoothly faired0.07
Windshield- Sharp edged0.15
Open Cockpit0.50
Table 6: Drag in different part of AircraftDrag due to lift
factor: K = 1/Ae {Here, A = Aspect Ratio, e = Oswalds Span
efficiency Factor= 0.75}Maximum Lift (Clean): CL max = 0.9 CL max
Cos (sweep) {CL max = for Airfoil}It is very difficult to calculate
the CL max accurately when the flap are used. But a quick
approximation can be made by the following formula.CL max = CL max
clean + 0.9 CL max {
Figure 16: S_flapped areaLift IncreaseDelta CL max
Plan & Split Flaps0.9
Slotted Flaps1.3
Fowler Flap1.3
Table 7: Lift enhanced data for flapsNow, propeller efficiency
can be calculated from the two key parameter Advance Ratio and
Power Efficiency.Advance Ratio: J = V/nD {Here, n = Propeller
rotational rate (rev/sec), D = Propeller Diameter, V=
Velocity}Power Efficiency: Cp = 550 bhp/n3D5 {Here,= Air Density
(altitude), bhp= Engine brake horsepower at that altitude}.Range: R
= , Where, L/D =Stability AnalysisIt is being rarely deserve by a
designer that aircraft is neutrally stable, rather the CG is
forward of the Neutral Point Xnp is deserved. In fact the distance
between the CG and Xnp will identify how stable the aircraft is?Xnp
= Where: , =Location of the quarter chord of the wing MAC. =
Location of the quarter chord of horizontal tail MACWfuselage=
Maximum fuselage width.Wfuselage = Fuselage length.Lratio =
Distance from front of fuselage to 25% of wing root
chord/LfuselageW :KTail Term= = 0.6 is probably reasonable (0.7 for
a T-tail), Its Effect of the wing turning the airflow before it
reach the tail.Detail DrawingAfter having done all the analysis,
detail drawing supposed to produce, for different components of
aircraft including the joints, nut, volts and every necessary
parts, which is being needed during manufacturing. If the
Ultralight aircraft is made for commercial purpose, in that case
detail drawing should be needed, it can also help to certifying the
aircraft. But for homebuilding it is not a must be fact, moreover
it would be good if the Homebuilders do have that.
Figure 17: Detail Drawing of AircraftPrototypeNow this the part
where the designer can give birth of his creation. Its perhaps the
dream comes into real existence. It is necessary to make the
prototype to check the all calculated vales in real life
environment. It supposed to make in such scale that is being
examined with the particular wind tunnel. Other than there is
another easy method can be followed by printing the CAD design with
3D printer. Now the prototype can judge by the wind tunnel
experiment and if there is huge tolerance, some parts of the design
supposed to be redone. Results & EvaluationThe main idea behind
this part of the project is to testify the whole methods, which has
been discussed above is correct or wrong. There is some calculation
is being made depending upon the above discussed equations. And
which is displayed in the following table for testification.
Aircraft
TypeMW.PWESARVSKm/hVSFFKm/hVMaxKm/hRCm/sSTOG{m}SLG{M}CLmaxCLmaxFF
P92 ECHO 80Ah27570.62334.4313.206.5571612105.51101001.401.90
P96 GOLF 80AL27570.62361.8412.205.7871612254.51101001.522.06
REMOS G-3CH27570.62366.6512.047.9875632206.5801401.381.95
DF 2000AH27470.62367.8812.008.3366562155.5110 1001.792.48
YUMA (STOL)AH2766 0.63328.4613.447.07 55501756.0 40
552.302.78
SAVANNAHAH2668 0.60343.8112.846.28 50 45 1606.0 50502.913.59
ZENAIR CH 701AH2580 0.58387.2411.405.9053 481537.05050
2.923.56
AMIGO !AL2806 0.64339.5813.005.2474 642506.580 1001.311.75
SLEPCEVSTORCH Mk4AH2649 0.60275.9116.006.765246 1554.550
502.162.76
SKY ARROW450TCH2825 0.64326.7613.516.9670 61192
5.1120801.411.86
Allegro 2000AH2727 0.62387.2411.4010.237363220
5.01501001.542.06
SINUS 912MotoalianteCH2786 0.63360.0712.2618.286663
2206.5881001.751.92
AVIO J-JabiruCH2649 0.60474.179.319.497464
2166.06.01601.832.45
EV-97 EUROSTAR Model2001AL2570 0.58448.639.846.6775 65225
5.5125901.692.25
JET FOX 97AH2845 0.64301.9514.626.547060175
6.01001201.301.77
TL 96 StarAL2747 0.62364.8312.106.8780 63250
6.0901001.211.94
Table 8: Design parameters of different AircraftHere,In Table: 8
{Weights, sizes and performances at sea level of the analyzed
aircraft (M. Material: A -aluminum alloy, C composite; W.P. Wing
Position: h high, l low}
AR: Wing Aspect Ratio.CLmax : Maximum Lift coefficient of the
aircraft with re-traced flaps.CLmaxFF : Maximum Lift coefficient of
the aircraft with full flaps.CLmaxL: Maximum Landing Lift
coefficient of the aircraft.CLmaxTO: Maximum Takeoff Lift
coefficient of the aircraft.RC: Maximum Rate of Climb.S : Wing
area.SLG: Landing Ground run. STOG: Take Off Ground run.Vmax:
Maximum level speed.Vmin: Minimum level speed.Vs: Stalling speed
flaps up.,VsFF: Stalling speed flaps down.WE: Empty Weight.WTO:
Maximum Takeoff Weight.
New-fangled Aircraft design:Empty weight1922 NTail:
Max T-O and landing weight3090 NHorizontal Stabilizer
{m}1.40
Max wing loading227 N/m2Vertical Stabilizer {m}0.75
Max power loading74 N/kWHorizontal:
Aircraft Length {m}4.56Span1.96
Aircraft Height {m}0.95Root chord0.50
CLmaxTO1.71Tip chord0.50
CLmax2.18Aspect ratio3.92
Wing:Vertical:
Wing Span6.80Span1.02
Root chord {m}0.98Root chord0.60
Tip chord {m}0.98Tip chord0.42
Aspect ratio (AR)6.93Aspect ratio2.00
Incidence {deg}2.00Leading edge sweep angle {deg}22.20
Wing Area {m2}9.52Trailing edge sweep angle {deg}13.00
Ailerons {m2}0.85Engine and Propeller:
Flap {m2}2Engine Power56 hp
Propeller Blade Number3
Propeller Diameter {m}1.16
Table 9: Design parameters of certain aircraftPerformance
Max speed {km/h}136Max rate of climb {m/s}4.70
Cruising speed {km/h}116Take off run {m}55
Stall speed {km/h}flaps up46Landing run {m}50
Table 10: Aircraft performance dataIf the above [Table: 9 and
Table: 10] data is being compared with the Table: data, then it is
being observed that the output data of Table: 9 and Table: 10 is
quite similar to the Table: 8 data (Those aircrafts are
successfully made their flight and proved their performance).
Hence, from the above discussion it is being justified that the
discussed Ultralight Aircraft Design Philosophical view can be
implemented for new conventional Ultralight Aircraft design
Conclusion and Outlook:The growing interest of Ultralight Aircraft
design and manufacturing in whole world is behind the inspiration
and drive to do such project, which would help the inexperienced
and new homebuilders to design their particular ultralight aircraft
in more abridged as well as cost effective way. The ultimate goal
of this project is to identify the key factors, which plays an
important role in ultralight aircraft design is being identified
and placed them in factual order. However, many of the equations,
used in this investigation in one hand is not truly accurate but on
the other hand they are quite impressive to predict the value. To
optimize the value further exploration is indispensable. The future
work should also aim at the understanding of the potential
equations and their optimization. Summary The main unprejudiced
part of this investigation is to find out the simplified aircraft
design methods relevant with the ultralight aircraft design
philosophy. This investigation includes an elaborate description of
the procedures, protocols and moreover deeper most findings of
design philosophy of ultralight aircraft as well as it is an
outmost learning guideline for ultralight aircraft design. From
this research its tried to figure out and prove that which come
after what? Its mean that clarifying the sequential order of
Ultralight Aircraft design.Many things about the aircraft design is
not described elaborately here but to provide a mature overview of
aircraft design as well as develop the particular notion, this
report would have been a landmark for an incompetent designers and
learners.To design a successful aircraft it is necessary to select
the starting point and follow the steps, which is being tried to
demonstrate here in perfect manner. To work as a professional
designer, acquaintance and know-how is needed, so in that case
Ultralight Aircraft Design Philosophy would perform as a platform
to ride in the design team of an ultramodern aircraft.
ReferencesBooks Donald R. Carwford A Practical Guide to Airplane
Performance and Design 1981. Mohammad H. Sadraey Aircraft Design A
Systems Engineering Approach 2013 John D. Anderson Jr. Aircraft
Performance and Design 1999. Michael Chun-Yung Niu
Airframe-Stuctural-Design 1988. Dr. Jan Roskom Airplane Design 1-8
vol 1985. Dr. Jan Roskom, Dr. Chuan Tau Edward Lan Airplane
Aerodynamics and Performance 1997. Martin Hollmann Composite
Aircraft Design 1983. Andy J. Keane and Prasanth B. Nair
Computational Approach on aircraft Design 2005. L. Pazmany Light
airplane design 1963. Egbert Torenbeek Synthesis of Subsonic
Airplane Design 1982. Friedrich Mller Flugzeugentwurf 2003 (German
writer).8.2. Reports, Article, Thesis and Individual Papers: AGARD
Report 783 Engineering Methods in Aerodynamic Analysis and Design
of Aircraft January, 1992. Chris Hainz Aircraft Design Made Easy
EAA Experimenter Magazine, November 2002. Chris Hainz Light
Aircraft raw Materials EAA Experimenter Magazine, March 1986.
Michael Case, Jeff Haack, Moon Chang Kim, Mandar Kulkarni, Darin
Mohr, Helmi Temimi Mathematical Techniques for Pre-conceptual
Design University of Minnesota, August 17, 2007. Assoc. Prof. Dr.
Serkan zgen, METU, Dept. Aerospace Eng The philosophy of airplane
design Fall 2009-2010. Omran Al-Shamma and Dr. Rashid Ali Aircraft
weight estimation in interactive design process University of
Hertfordshire. Alex Paterson Aspects of Aircraft Design that
Enhance Safety an aviation safety article, 21 September 2007. F. H.
Darwish, G. M. Atmeh, Z. F. Design Analysis and Modeling of a
General Aviation Aircraft Hasan Jordan Journal of Mechanical and
Industrial Engineering, April 2012. D. P. Coiro, A. de Marco, F.
Nicolosi, N. Genito, S. Figliolia Design of a Low-Cost Easy-to-Fly
(STOL) Ultralight Aircraft in Composite Material Czech Technical
University in Prague, Acta Polytechnica Vol. 45 No. 4/2005. Design
of 4 seat general aviation, electric aircraft, San Jose State
University, California, USA, 2012. Dr. Randal Allen Airplane
performance and Design 2012. EASA Certification Specifications for
Very Light Aeroplanes CS-VLA 5 March 2009.8.3. Electronic
Publications: http://articles.sae.org/12585/
http://articles.sae.org/12473/
http://www.sciencedirect.com/science/article/pii/S1369886901000052
http://www.aerostudents.com/files/aircraftPerformance2/aircraftPerformance2FullVersion.pdf8.4.
Software XFLR5 Homebuilt Aircraft Design Spreadsheet Martin
Hollmann Software Appendix A - Air properties and conversions
Appendix B - Unit Conversions for Design
Appendix D - Aircraft Materials Density
Appendix E Relevant websites address
http://aerospace.illinois.edu/m-selig/ads/coord_database.html
http://www.xflr5.com/xflr5.htm
http://www.nasascale.org/howtos/mac-calculator.htm
http://fwcg.3dzone.dk/ http://adamone.rchomepage.com/cg_calc.htm
http://chrusion.com/BJ7/SuperCalc7.html
http://web.mit.edu/drela/Public/web/avl/
http://www.mathsisfun.com/geometry/dihedral-angles.html