Ultra Light Aircraft Design Philosopy

Project Report of Elective Studies

Ultralight Aircraft Design Philosophy Submitted by:

Khandoker Raisul Azad

Registration Number:

MAV-1310588003

Academic Supervisor:

Eva Windbacher-Schwager

Date of Submission:

30/01/2014

Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

Page I

Preface Different 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 can’t 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 ladder’s 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.


Page II

Contents 1.

Introduction .......................................................................................................................................... 6

2.

Methods ................................................................................................................................................ 6 2.1.

Identify the Key factors ................................................................................................................. 6

2.2.

Historical Data ............................................................................................................................... 6

2.3.

Literature Review .......................................................................................................................... 7

2.4.

Consultations and Discussions ...................................................................................................... 7

2.5.

Technological Resources ............................................................................................................... 7

2.5.1.

Software ................................................................................................................................ 7

3.

Design Philosophy ................................................................................................................................. 8

4.

Design Phase ......................................................................................................................................... 8 4.1.

Requirements ................................................................................................................................ 9

4.2.

Initial Drawing ............................................................................................................................. 10

4.3.

Weight Calculation ...................................................................................................................... 10

4.3.1.

Statistical Rule ..................................................................................................................... 10

4.3.2.

Newton's Weight Equation ................................................................................................. 10

4.3.3.

Per Square Foot Method ..................................................................................................... 12

4.4.

Engine and Propeller Selection ................................................................................................... 12

4.4.1.

Propeller Selection .............................................................................................................. 12

4.5.

Airfoil Selection ........................................................................................................................... 13

4.6.

Wing Design ................................................................................................................................ 15

4.6.1.

Tapper Ratio: coefficient ..................................................................................................... 16

4.6.2.

Wing Shape ......................................................................................................................... 17

4.6.3.

Dihedral Angel ..................................................................................................................... 18

4.7.

Tail Geometry.............................................................................................................................. 19

4.8.

Fuselage Sizing ............................................................................................................................ 20

4.9.

Landing Gear design .................................................................................................................... 21

4.10.

Aircraft’s Fragment Placement ............................................................................................... 26

4.10.1.

Engine Location ................................................................................................................... 26

4.10.2.

Tail Placement ..................................................................................................................... 27

4.10.3.

Wing Hinge Point ................................................................................................................ 27

4.10.4.

Landing Gear Location ........................................................................................................ 29


Page 3

4.11.

Control Surface Design ............................................................................................................ 30

4.12.

Further Weight Calculation ..................................................................................................... 31

4.13.

Performance Analysis.............................................................................................................. 33

4.14.

Stability Analysis...................................................................................................................... 35

4.15.

Detail Drawing......................................................................................................................... 35

4.16.

Prototype ................................................................................................................................ 36

5.

Results & Evaluation ........................................................................................................................... 36

6.

Conclusion and Outlook: ..................................................................................................................... 38

7.

Summary ............................................................................................................................................. 39

8.

References .......................................................................................................................................... 39 8.1.

Books ........................................................................................................................................... 39

8.2.

Reports, Article, Thesis and Individual Papers: ........................................................................... 40

8.3.

Electronic Publications: ............................................................................................................... 40

8.4.

Software ...................................................................................................................................... 40

9.

Appendix A - Air properties and conversions ..................................................................................... 41

10.

Appendix B - Unit Conversions for Design ...................................................................................... 42

11.

Appendix D - Aircraft Materials Density ......................................................................................... 43

12.

Appendix E – Relevant websites address ........................................................................................ 43

List of Figures Figure 1: Ultralight Aircraft Design Phase ---------------------------------------------------------------------------------- 9 Figure 4: Airfoil Selection flowchart ---------------------------------------------------------------------------------------- 14 Figure 5: Wing Design Process Flowchart --------------------------------------------------------------------------------- 15 Figure 6: Wing Tapper Ratio -------------------------------------------------------------------------------------------------- 17 Figure 7: Mean Aerodynamic Chord ---------------------------------------------------------------------------------------- 18 Figure 8: Landing Gear Design Flowchart---------------------------------------------------------------------------------- 22 Figure 9: Landing Gear Primary Parameters ------------------------------------------------------------------------------ 23 Figure 10: Wheel Load Geometry ------------------------------------------------------------------------------------------- 24 Figure 11: Landing Gear CG Location --------------------------------------------------------------------------------------- 25 Figure 12: Top and Side view of Aileron ----------------------------------------------------------------------------------- 30 Figure 3: Wing structural Analysis ------------------------------------------------------------------------------------------- 33 Figure 13: S_flapped area ----------------------------------------------------------------------------------------------------- 34 Figure 14: Aircraft CG Calculation ------------------------------------------------------------------------------------------- 29 Figure 15: Detail Drawing of Aircraft --------------------------------------------------------------------------------------- 36


Page 4

List of Tables Table 1: Materials weight in different part of Aircraft ----------------------------------------------------------------- 12 Table 3: Tail geometry Ratios------------------------------------------------------------------------------------------------- 20 Table 4: Different Aircraft Aileron data ------------------------------------------------------------------------------------ 31 Table 2: Empty weight fractions of different materials ---------------------------------------------------------------- 32 Table 5: C_fe Data --------------------------------------------------------------------------------------------------------------- 34 Table 6: Drag in different part of Aircraft --------------------------------------------------------------------------------- 34 Table 7: Lift enhanced data for flaps --------------------------------------------------------------------------------------- 35 Table 8: Design parameters of different Aircraft ------------------------------------------------------------------------ 37 Table 9:Design parameters of certain aircraft---------------------------------------------------------------------------- 38 Table 10: Aircraft performance data --------------------------------------------------------------------------------------- 38


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1. Introduction Aircraft 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 designer’s 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, it’s 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.

2. Methods Different methods had been followed to find out the correct philosophical view for ultra-light aircraft design which is being described as follows.

2.1.

Identify the Key factors

There 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.

2.2.

Historical Data

Aircraft 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.


Page 6

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).

2.3.

Literature Review

In 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.

2.4.

Consultations and Discussions

Some 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.

2.5.

Technological Resources

To 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.

2.5.1. Software There 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 Software


Page 7

XFLR5 is one of the software who can save the designer’s 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.

3. Design Philosophy For 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 good ergonomics 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.

4. Design Phase This 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.


Page 8

Ultralight Aircraft Fuzelage Sizing

Landing Gear Design •Fixed

Requairments •Use Certification Data

Tail Geometry •Horizontal Sizing •Vertical Sizing

Initial Drawing

Weight Calculations

Prototype

Control Surface Design •Aileron •Elivator •Radar •Flap

Detail Drawing

Wing Selection •C L Max Calculation (3D) •Inclination angel •Wing area •Wing Loading

Engine Selection

Aerofoil Selection

•Propeller Diameter •Weight Power Ratio

•C L Max Calculation (2D)

Aircraft Part Placement •Engine location •Tail location • Wing hinge point •Landing gear location

Re-calculation of Weight

Stablity Analysis

N0

Performance Analysis

Figure 1: Ultralight Aircraft Design Phase

4.1.

Requirements

It 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 Ultralight 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


Page 9

    

Cruise Height Engine Lycoming Wing Area Wing Chord Wing Span

4.2.

Initial Drawing

To 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.

4.3.

Weight Calculation

This 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.

4.3.1. Statistical Rule But 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.

4.3.2. Newton's Weight Equation The following rule can be followed for the weight calculation: Using Newton's weight equation:


Page 10

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 unique density. Density “d” is defined to be the mass divided by the volume v: So, d = m/v { m= Mass & v = Volume} or, m = r × v To find out the Volume-“v” the following formula can be followed.

Cube

side3

Rectangular Prism side1 × side2 × side3 Sphere

(4/3) × pi × radius3

Ellipsoid

(4/3) × pi × radius1 × radius2 × radius3

Cylinder

pi × 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


Page 11

      

Propulsion System Instruments and Navigation Electrical System Electronics Crew Payload Fuel

4.3.3. Per Square Foot Method There 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 Estimation Metal Fiberglass Carbon fiber Fabric

Wing lb/sq-ft 1.1 to 2.0 1.6 to 2.2 1.2 to 2.0 1.0 to 2.0

Horizontal Tail lb/sq-ft 0.9 to 2.0 0.9 to 2.0 0.9 to 2.0 0.8 to 1.5

Vertical Tail lb/sq-ft 0.9 to 2.0 0.9 to 2.0 0.9 to 2.0 0.8 to 1.5

Fuselage lb/sq-ft 1.2 to 1.4 1.2 to 1.4 0.7 to 1.2 1.4 to 1.8

Table 1: Materials weight in different part of Aircraft

4.4.

Engine and Propeller Selection

If 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 can’t do anything without a propeller, so it’s 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-0.75 By 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}

4.4.1. Propeller Selection The 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) Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

Page 12

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=√

4.5.

{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.


Page 13

Figure 2: Airfoil Selection flowchart

There 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 doesn’t create unwanted drag. Calculate the lift coefficient during cruise by following the below equation, then find the angle of attack (α) that lift coefficient. Wing Loading: W/S=qCL, Where, q= ½ ρVmax2 And, CL= CLmax So, cruise lift coefficient: CL-cruise= Angel of attack: α = CL-Cruise{10+18/A}+αzerolift Or, If the wing is Swift: α = CL-Cruise

+ αzerolift

Here, A= Wing Aspect Ratio. αzerolift 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 that’s 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.


Page 14

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.

4.6.

Wing Design

Wing design is quite challenging part of aircraft design, this part and airfoil selection part’s 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 Flowchart


Page 15

For 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 = ½ ρV2S CL Or, W = ½ ρV2S CL Or, CL = ½ ρV2S W For a typical ultralight aircraft, CL-max will be about 1.4 without using flaps.

4.6.1. Tapper Ratio: coefficient Tapper Ratio (λ) 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).


Page 16

Figure 4: Wing Tapper Ratio

4.6.2. Wing Shape Now, by using the following formula the wing shape can be visualized. Wing Span: b = √ Root Chord: Croot = 2S/b (1+λ) {λ = Tapper ratio} Tip Chord: Ctip = λ Croot Now 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”.


Page 17



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 Chord

The 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. λ = Taper Ratio = (Tip Chord ÷ Root Chord) MAC: ̅ = Croot x 2/3 x ((1 + λ + λ2 ) ÷ ( 1 + λ ))

4.6.3. Dihedral Angel The 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.


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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 between−15 to +10 deg.

4.7.

Tail Geometry

Aircrafts 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: Vertical Tail:

= CHT ( ̅ = CVT (

/LHT /LVT

Here,

LHT = Tail moment arm length (for Horizontal tail) LVT = Tail moment arm length (for Vertical tail) b wing = Wing Span ̅ wing = 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). Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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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: Vertical Stabilizer Root Chord:

{Here,

}

{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 it’s 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 Geometry

Horizontal Aspect Ratio Taper Ratio 3 to 5 0.3 to 0.6 3 to 5 0.3 to 0.6 6 to 10 0.3 to 0.5

Conventional T-tail Sailplane

Vertical Aspect Ratio Taper Ratio 1.3 to 2.0 0.3 to 0.6 0.7 to 1.2 0.6 to 1.0 1.5 to 2.0 0.4 to 0.6

Table 2: Tail geometry Ratios

Elevators 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.

4.8.

Fuselage Sizing

How big the fuselage would be? The answer of the question depends upon the designer’s 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

0.23

Of 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 Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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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.

4.9.

Landing Gear design

Landing 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


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Figure 7: Landing Gear Design Flowchart


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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


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     

Aircraft structural integrity Ground lateral stability Low cost Low weight Maintainability Manufacturability

Wheel 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. ½ 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 Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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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 Location

For 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:


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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.

4.10. Aircraft’s Fragment Placement After 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.

4.10.1. Engine Location In 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 CG

Considering 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. Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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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.

4.10.2. Tail Placement The 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.

4.10.3. Wing Hinge Point Now 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 location

Estimated CG:

⁄

Here,


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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 { value]} behind the nose of the aircraft.

wing

= [calculated

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.


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Ref: http://www.grc.nasa.gov/WWW/k-12/airplane/acg.html

Figure 12: Different objects CG calculation

4.10.4. Landing Gear Location As 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.


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Figure 13: Landing gears position

4.11. Control Surface Design As 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. 2. 3. 4.

Aileron plan form area (Sa); Aileron Chord/Span (Ca/ba); Maximum up and down aileron deflection (+ or - δAmax); and 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 δAmax= 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 Aileron

Therefore the following statistical data of different aircraft would develop some idea on aileron designing. Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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Table 3: Different Aircraft Aileron data

4.12. Further Weight Calculation Otherwise 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 = 1/0.75πA = 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. Ultralight Aircraft Design Philosophy (2014, Graz, Austria)

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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/550ηpL/D Here, the 0.975 term is used an approximate allowance for additional fuel used during takeoff, climb, descend and landing. The value of “e=2.7183”approximately. “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 doesn’t 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 Type Metal or Wood Design Composite Design

Single Engine 1.19 1.15

Twin Engine 1.40 1.35

Table 4: Empty weight fractions of different materials

The 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/W0 It is indeed be important to calculate the thickness of the structural parts required to safety withstand the expected loads, including a factor of safety.


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Figure 15: Wing structural Analysis

In 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 =

4.13. Performance Analysis It 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, ρ = air density at sea level, W = Weight, S = Wing Area}

Takeoff Parameter: T.O.P = 1.21 Rate

of

Climb:

Vv=V{ ⁄

⁄

⁄

⁄

{Here, HP = Horse Power, W = Weight, S = Wing Area} (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).


Page 33

Total Drag: DTotal = Here, CD0= Cfe

S(CD0 + KCL2)

[Cfe= Equivalent skin friction coefficient, Sref= Wing reference area.

Cfe Average Metal Design Smooth Composite

Single Engine 0.0058 0.0050

Twin Engine 0.0048 0.0045

Sail Plane 0.0038 0.0030

Table 5: C_fe Data

Drag Area Exposed Wheel & Tire nd 2 wheel in Tandem Streamlined wheel and tire Wheel & Tire “pants” Round Strut Streamlined Strut Flat Spring Gear Leg Fork or Irregular Fitting Speed brake- Fuselage Speed brake- Wing Windshield- Smoothly faired Windshield- Sharp edged Open Cockpit

Drag/Dynamic Pressure (D/q) per unit frontal area 0.25 0.15 0.18 0.13 0.30 0.05 1.40 1.0 to 1.4 1.00 1.60 0.07 0.15 0.50

Table 6: Drag in different part of Aircraft

Drag due to lift factor: K = 1/πAe {Here, A = Aspect Ratio, e = Oswald’s 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 area

Lift Increase Plan & Split Flaps Slotted Flaps Fowler Flap

Delta CL max 0.9 1.3 1.3


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Table 7: Lift enhanced data for flaps

Now, 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/ρn3D5 {Here,ρ= Air Density (altitude), bhp= Engine brake horsepower at that altitude}. Range: R =

, Where, L/D = ⁄

(

)

4.14. Stability Analysis It 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 MAC Wfuselage= Maximum fuselage width. Wfuselage = Fuselage length. Lratio = Distance from front of fuselage to 25% of wing root chord/Lfuselage W

:

KTail Term= ” = 0.6 is probably reasonable (0.7 for a T- il I ’ Effect of the wing turning the airflow before it reach the tail.

4.15. Detail Drawing After 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.


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Figure 17: Detail Drawing of Aircraft

4.16. Prototype Now this the part where the designer can give birth of his creation. It’s 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.

5. Results & Evaluation The 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.


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Aircraft Type

M

WE

A A C A A A A A A

W .P h L H H H H H L H

S

AR

P92 ECHO 80 P96 GOLF 80 REMOS G-3 DF 2000 YUMA (STOL) SAVANNAH ZENAIR CH 701 AMIGO ! SLEPCEV STORCH Mk4 SKY ARROW 450T Allegro 2000 SINUS 912 Motoaliante AVIO J-Jabiru EV-97 EURO STAR Model 2001 JET FOX 97 TL 96 Star

VSFF Km/h 61 61 63 56 50 45 48 64 46

VMax Km/h 210 225 220 215 175 160 153 250 155

RC m/s 5.5 4.5 6.5 5.5 6.0 6.0 7.0 6.5 4.5

STOG {m} 110 110 80 110 40 50 50 80 50

SLG {M} 100 100 140 100 55 50 50 100 50

CLmax

CLmaxFF

6.55 5.78 7.98 8.33 7.07 6.28 5.90 5.24 6.76

VS Km/h 71 71 75 66 55 50 53 74 52

2757 2757 2757 2747 2766 2668 2580 2806 2649

0.62 0.62 0.62 0.62 0.63 0.60 0.58 0.64 0.60

334.43 361.84 366.65 367.88 328.46 343.81 387.24 339.58 275.91

13.20 12.20 12.04 12.00 13.44 12.84 11.40 13.00 16.00

1.40 1.52 1.38 1.79 2.30 2.91 2.92 1.31 2.16

1.90 2.06 1.95 2.48 2.78 3.59 3.56 1.75 2.76

C

H

2825

0.64

326.76

13.51

6.96

70

61

192

5.1

120

80

1.41

1.86

A

H

2727

0.62

387.24

11.40

73

63

220

5.0

150

100

1.54

2.06

0.63

360.07

12.26

66

63

220

6.5

88

100

1.75

1.92

2649 2570

0.60 0.58

474.17 448.63

9.31 9.84

10.2 3 18.2 8 9.49 6.67

C

H

2786

C A

H L

74 75

64 65

216 225

6.0 5.5

6.0 125

160 90

1.83 1.69

2.45 2.25

A A

H L

2845 2747

0.64 301.95 14.62 6.54 70 60 0.62 364.83 12.10 6.87 80 63 Table 8: Design parameters of different Aircraft

175 250

6.0 6.0

100 90

120 100

1.30 1.21

1.77 1.94

⁄

⁄

Here, 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 CLmax CLmaxFF CLmaxL CLmaxTO RC S SLG STOG Vmax Vmin Vs

: Wing Aspect Ratio. : Maximum Lift coefficient of the aircraft with re-traced flaps. : Maximum Lift coefficient of the aircraft with full flaps. : Maximum Landing Lift coefficient of the aircraft. : Maximum Takeoff Lift coefficient of the aircraft. : Maximum Rate of Climb. : Wing area. : Landing Ground run. : Take Off Ground run. : Maximum level speed. : Minimum level speed. : Stalling speed flaps up.

,VsFF WE WTO

: Stalling speed flaps down. : Empty Weight. : Maximum Takeoff Weight.


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New-fangled Aircraft design: Empty weight Max T-O and landing weight Max wing loading Max power loading Aircraft Length {m} Aircraft Height {m} CLmaxTO CLmax Wing: Wing Span Root chord {m} Tip chord {m} Aspect ratio (AR) Incidence {deg} Wing Area {m2} Ailerons {m2} Flap {m2}

1922 N 3090 N 227 N/m2 74 N/kW 4.56 0.95 1.71 2.18 6.80 0.98 0.98 6.93 2.00 9.52 0.85 2

Tail: Horizontal Stabilizer {m} Vertical Stabilizer {m} Horizontal: Span Root chord Tip chord Aspect ratio Vertical: Span Root chord Tip chord Aspect ratio Leading edge sweep angle {deg} Trailing edge sweep angle {deg} Engine and Propeller: Engine Power Propeller Blade Number Propeller Diameter {m}

1.40 0.75 1.96 0.50 0.50 3.92 1.02 0.60 0.42 2.00 22.20 13.00 56 hp 3 1.16

Table 9: Design parameters of certain aircraft

Max speed {km/h}

Performance 136 Max rate of climb {m/s}

4.70

Cruising speed {km/h} Stall speed {km/h}flaps up

116 46

55 50

Take off run {m} Landing run {m}

Table 10: Aircraft performance data

If 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

6. 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.


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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.

7. 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 it’s tried to figure out and prove that which come after what? It’s 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.

8. References 8.1.           

Books

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 Müller “Flugzeugentwurf” 2003 (German writer).


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8.2.    

    

  

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 Özgen, 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.pdf

8.4.   

Reports, Article, Thesis and Individual Papers:

Software

XFLR5 Homebuilt Aircraft Design Spreadsheet Martin Hollmann Software


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9. Appendix A - Air properties and conversions


Page 41

10.

Appendix B - Unit Conversions for Design


Page 42

11.

Appendix D - Aircraft Materials Density

12.

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/


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

http://www.mathsisfun.com/geometry/dihedral-angles.html


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Ultra Light Aircraft Design Philosopy

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