Wing and Tail Design 1 1 PDF

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Wing and Tail Design DESIGN DECISIONS AND CALCULATIONS FOR THE WING AND TAIL. ALAA ADEL

201227908

HARRY THOMAS

201157413

JACK TULLEY JOE PIGOTT

201187517 201188930

SUNGJIN PARK

201144070

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Abstract As the conceptual design and preliminary design are completed and some parameters have been derived, the detailed design can be done. In this designing phase, the main wing aerofoil is determined by the consideration of some standards such as maximum lift coefficient, design lift coefficient and lowest minimum drag coefficient. The other 18 main wing parameters (aspect ratio, taper ratio, tip chord, root chord, mean aerodynamic chord, span, twist angle, sweep angle, dihedral angle, incidence, high-lifting wing devices, aileron) also have been calculated. Some figures such as aspect ratio were taken from the model aeroplane (A400M) and it was determined as 7.986. Some figures such as taper ratio and dihedral angle are used the same as the tail wing as in the main wing. In terms of tail wing design, it should consider three requirements: trim, stability and control. 26 horizontal and vertical tail parameters have been determined throughout the detail design stage. An Aerofoil also has been selected to satisfy the conditions and a Horizontal tail wing span has been determined as 18.08m. Throughout wing detail design stage, all the parameters related with wing and tail wing have been determined.

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Contents Abstract………………………………………………………………………………………………………………..1 Table of Contents………………………………………………………………………………………………….2 1. Wing Design………………………………………………………………………………………………………3 2. Tail Design…………………………………………………………………………………………………………5 3. Conclusions……………………………………………………………………………………………………….7 4. Recommendations…………………………………………………………………………………………….7 6. References…………………………………………………………………………………………………………8 7. Appendices………………………………………………………………………………………………………..9

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1. Wing Design After the conceptual and preliminary design phases where done, the next step is going to be detail design. The phase where most of the major parts of the aircraft are going to be designed in a detailed way such as the wing, fuselage, horizontal and vertical tails. The major parts mentioned and others are going to be designed individually, then comes the process of getting everything together later on. According to M. Sadraey, as mentioned in his book “Aircraft design: A system engineering approach”, wings are the most important aircraft component. The wing’s primary function is to generate aerodynamic lift, as anything cannot be taken for nothing, they produce 2 other aerodynamic forces- aerodynamic drag and nose pitching moment. Ideally lift should be maximized while drag and pitching moment should be minimized. Wings in design phase goes through a process that consists of 18 factors all must be fulfilled. The parameters as recognized in “Aircraft design: A system engineering approach” are: Wing reference area (calculated in the preliminary phase), Number of wings, Vertical position relative to the fuselage, Horizontal position relative to the fuselage, Aspect ratio, Aerofoil, Taper ratio, Tip chord, Root chord, mean aerodynamic chord, Span, Twist angle, Sweep angle, Dihedral/ anhedral angle, Setting angle, High-lifting devices, Aileron, Other wing accessories Number of wings: All modern aircrafts use monoplanes as conventional number of wings. Based on modern technology a single wing was chosen to prevent the disadvantages related to using any number of wings other than monoplane. The disadvantages can be represented in a form of more weight, less lift and pilot visibility limitations. Wing vertical location: The vertical location of the wing is one of the parameters that must be determined at early stages of the wing design. The wing vertical location has an impact on other aircraft components as tail design and landing gears. Following conventional design lines, most cargo and general aviation aircrafts have high wings, working simultaneously with the primary operational requirements and some influencing factors such as stability and producibility. High-wing was the choice. Before the final decision was made, pros and cos were well studied. Some major advantages of high wings are that they are easier in terms of the loading and unloading process, increases lateral stability and increases downwards visibility, but penalty has to be paid in form of the disadvantages. High-wing location decreases lower ground effect, increases the fuselage weight and decreases upwards visibility. Aerofoil selection: By definition; it is a special wing cross sectional, Aerofoils are 2D (Sydney,2013). Aerofoil is the third- out of 18- in the design list and the second most important parameter, it comes after wing planform area. The importance came from the task it should fulfil. The Aerofoil is responsible for the pressure distribution over the 2 wing surfaces leading to lift creation. They can be designed or selected, designing an aerofoil from scratch is a time-consuming process and is more expensive than being selected. Aerofoils have major geometric features that will always be there, leading and trailing edges, meanline, chord line, camber line- the camber lines are usually positive- and the aerofoil thickness. Any change even a small one in any of the mentioned parameters will end up with a different shape. The aerodynamic centre of gravity- is where the aerodynamic momentum is independent of the angle of attack- and aerodynamic centre of pressure- the point where the resultant aerodynamic force and the aerodynamic momentum is zero- are points located on the aerofoil (week 7). For A400M Atlas the aerofoil used was 654-221. Wing incidence: Wing incidence; is the angle between the fuselage centreline and the wing chord at its root. Wing incidence must fulfil basic design requirements, it must be able to generate lift during straight and level flight (cursing), produce as low a drag as possible and allow angle of attack to increase within safe limits during take-off. The angle can be selected to be constant throughout the flighting operations or be variable. For this model the wing incidence is constant. Choosing constant wing incidence, the wing setting must be found for manufacturing purpose. Aspect ratio Referring to the definition, it is the ratio between the wing span b and the wing C 3|Pa ge

AR=

b = 7.986 m^2 C

Aspect ratio has a direct impact on some of the wing properties and the aircraft in general as performance, stability, control, cost and the manufacturing process. For wing design in process, the aspect ratio is that it is a mid-aspect ratio wing. As the decision was being made advantages and disadvantages were taking in consideration. Taper ratio It is the ratio between the tip chord and the root chord (Sydney,2013). Λ=

c tip =¿ c root

It varies between 0 and 1. For the A400M atlas, the taper ratio shape is trapezoid. By using this shape- not rectangular or delta- aerodynamic inefficiency was avoided, at the mean time some advantages were lost as well. The significance of lift and load distributions The distribution of wing non-dimensional lift per unit span along the wing is referred to as lift distribution (Sydney,2013). Each point of the in produces lift but they are not all the same. Lift production decreases near the wing tip as the pressure difference decreases to the point where the difference is zero-wing tip. Load distribution is the variation of lift coefficient times sectional chord (Sydney,2013). Lift distribution is mostly used in aerodynamic calculations while primary use of load distribution is in structural design. Sweep angle There are 2 types of sweep angles, Leading edge sweep angle and trailing edge sweep angle. The main reasons behind including sweep angle are; it improves the wing aerodynamic features by delaying the compressibility effects, improves lateral stability and longitudinal and directional stability, works on improving centre of gravity and increases pilot view. For this design the wing is positive leading-edge sweep angle. Twist angle For the A400M, the wing tip is lower incidence compared than the wing root, the wing has negative twist. This indicate that the angle of attack is lower at the tip root relative to the wing root, it decreases along the wing span. The main reasons behind angle of twist application are avoiding tip stall before root stall and modification of the lift distribution to an elliptical one (Sydney,2013). Dihedral angle The angle between the chord line of a wing with xy plane (Sydney,2013). Wing dihedral is mainly used to improve lateral stability where the aircraft has the tendency to return to its original trim level-wing flight conditions. A400M has negative dihedral or in other words anhidral angle where the tip of the wing is lower than the wing root. High-lift device HLDs are used in take-offs and landings where the airspeed is lower relative to cruise speed, based on that the wing must produce greater lift coefficient. HLDs deflection help in increasing lift coefficient maximum lift coefficient, drag coefficient and lift curve. It changes zero-lift angle of attack, stall angle and pitching moment coefficient. there are 2 main groups of HLDs leading edge high-lift device and trailing edge high-lift device(flap). Aileron A trailing edge look alike, the difference exists the way the deflection takes place, aileron deflects both up and down. It is located at the outboard portion at the left and the right section of a wing (book). Lateral control is applied on an aircraft through the differential motions. While designing an aileron 3 main factors has to be determined; aileron chord, aileron span and aileron deflection. It is mainly used to work on roll controllability.

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2. Tail Design The tail design has been decided as T tail configuration in previous stage. Tail wing has three functions: trim, stability, and control. According to Sadraey (2012), Tail wing parameters are decided through the consideration of those three requirements (trim, stability, and control). It is important to consider the requirements of trim for safe flight because in trimmed condition, the aircraft will not rotate about its centre of gravity but the aircraft will keep flying as straight level flying without any additional load on control column or desired circular motion. To be trimmed condition, the summations of all forces and moments should be zero.

l (¿ ¿ h , lv ) → Aerofoil section → Aspect ¿ ratio ( ARh , ARv ) → Taper ratio ( λh , λ v ) → Tip chord ( Ch , C v ¿ → Root chord ( Ch , C v ¿ → Mean

The process of the tail wing design is: Planform area ( S h , S v ¿

→ Tail arm

tip

tip

root

root

b ¿ b aerodynamic chord ( MAC h∨Ch , MAC v ∨C v ¿ → Span → Sweep angle ( Λh , Λv ¿ → Dihedral angle (¿¿ v) ¿ ¿ (Г h , Г v ) → Tail installation → Incidence ( i h , iv ¿ 2.1 Horizontal tail wing The following figures have been derived from previous stage or given as: MTOW ( m ¿ ¿=183236.42 kg , maximum fuselage diameter ( D fmax ¿=5.7 ( reference aeroplane− A 400 M ) , Cruise speed ( V c ¿=¿ 216.67m/s, angle of flap ( α f ¿=1deg (at cruise), main wing area

(S w )=¿ 225.1 m 2 , Wing chord ( C w ¿=5.31m , Aspect ratio (AR)=7.986, Taper ratio ( λ w ¿=¿ 0.4, wing incidence ( i w ¿=0.05 rad , wing twist ( α twist ¿=−1.1 deg , Leading edge sweep ( Λ¿ ¿=¿ 16 deg, dihedral angle ( Г w ¿ =0.09rad, main wing aerofoil: NACA 654 -221, coefficient of lift per angle of attack ( C L ¿ =5.8 1/rad α

Horizontal tail volume coefficient ( V HT ¿ is selected as 0.95 because general military cargo’s (A400M) tail volume coefficient known as 1.00 and typical twin turboprop horizontal tail volume coefficient value is averagely 0.9. Optimum tail moment arm ( lopt ) can be determined by minimising the aircraft drag.

√

l =l opt=K c

√

4 CS V H 4 × 5.31 × 225.1 × 0.95 =19.11m = 1.2 π ×5.7 π Df

Tail planform area can be derived by using equation

S h=

CS V HT 5.31 ×225.1 ×0.95 =61.4 m 2 = 19.11 l

Aircraft cruise lift coefficient:

C L=

2 W avg 2× 183236.42× 9.81 = =0.376 2 ρV c S 0.905 × ( 216.67 )2 ×225.1

(Where air density at 10000ft is 0.905 kg /m3 ¿

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V HT =

l Sh : CS

Aerofoil section pitching moment coefficient can be extracted from aerofoil ( Cm o ¿ wf

graph: value of Cm af : -0.025

(Sadraey, 2012).

Cm =C m o wf

af

ARco s 2 ( Λ) +0.01α t AR +2 cos( Λ)

=

7.986 × −0.025

( 12 + 21 cos 32 ) + 0.01×−1.1

=-0.030

7.986 +2 × cos ( 16 )

To use trim equation, h and ho should be calculated. It is assumed that the aerodynamic centre of the wing and fuselage combination is located at 23% of mean aerodynamic centre (MAC) and centre of gravity is located at 20% of MAC. ( X cg =20 %MAC ¿ h=0.2, and tail efficiency is assumed to be 0.98. By using the trim equation ( Cm o +C L ( h− ho )−ηh V HT C L h=0 ), the horizontal tail required lift coefficient at cruise can wf

be derived:

CL =

C m +C L( h−ho )

h

o wf

=

V HT

−0.030+0.376 ×(0.2−0.23) =−0.043 0.95

Aerofoil has been chosen by the consideration of properties: 1. Capacity of generating required lift with minimum drag and minimum pitching moment. 2. Symmetric. 3. The main wing should stall before tail wing. 4. Horizontal tail must be clean of compressibility effect. 5. Thinner than wing aerofoil to avoid compressibility. Among several aerofoil sections that satisfy these conditions, the aerofoil that has a low drag coefficient has been chosen as NACA 653−218 . The initial tail aspect ratio has been determined:

2 2 AR h= ARw = ×7.986 =5.324 3 3 Tail taper ratio is the same as wing taper ratio:

λh = λw =¿ 0.4 Tail taper sweep angle and tail dihedral angle are the same as wing

Λh =16 deg , Γ h=¿ 0.09rad Tail setting angle ( i h ¿ is determined by consideration of tail parameters and downwash. By based on the tail lift curve slope, tail angle of attack is determined. Tail generated lift coefficient can be calculated by lifting line theory. The tail generated lift coefficient should be equal to tail required lift coefficient. Tail incidence could adjust these figures to be equal.

Tail lift curve slope:

Cl

CL = α

1+

5.8 =¿ 5.8 = 4.31 1/rad 1+ π × 5.324

αh

Cl

αh

π AR h

Tail angle of attack in cruise: α h =

C L −0.043 = =−0.01deg CL 5.8 h

αh

Tail created lift coefficient with an angle of attack of -0.01deg is calculated as: -7.13 Tail is expected to generate C Lh of −0.043 but it generates -7.13. As it generates too much lift, tail lift coefficient should be reduced, so the tail angle of attack should be reduced. Therefore, the angle of attack for horizontal tail has been reduced as -0.63 deg. Downwash ( ε ): ε o+ 6|Pa ge

∂ε ∂α

ε o (downwash angle at zero angle of attack) =

2 CL πAR

=

w

2 ×0.376 π ×5.324

=0.045rad = 2.57deg

2 CL 2 ×5.8 ∂ε =0.69 deg / deg = ( downwashslope ) = πAR π ×5.324 ∂α αw

0.045+0.69 ×

3 =0.081 deg 57.3

Therefore, tail setting angle: α t=α f +i h−ε →

ih=α h−α f +ε → -0.63-1+0.081=-1.54deg

Other parameters are derived by the equations of: AR h=

bh Ch

Ch , λh = Ch

tip

1+ λh + λh2 2C ( , Ch = ) , 3 h 1+ λ h root

root

S h=b h C h The results are given as tail span ( bh )=18.08m, tail root chord ( Ch root ¿ =4.57m, tail tip chord ( Chtip ¿=1.83 m , horizontal tail mean aerodynamic chord (C h) =3.4m To examine the aircraft static longitudinal stability: Cm α=C Lα wf ( h− ho ) −C Lα ηh h

Cm =5.7( 0.2 −0.23 )−4.31 ×1× α

(

Sh l ∂ε ( −h)(1− ) S Ch ∂α

)

59.42 19.11 −0.2 ( 1−0.69 )=−2.081 /rad 225.1 3.4

( ηh=1 due ¿the T tail configuration ¿

2.2 Vertical tail Vertical tail volume coefficient ( V VT ) has been determined as 0.08 as military transport(A400M) and twin turboprop’s vertical tail volume coefficient is normally 0.08. Vertical tail moment arm

l (¿¿ v) is assumed as it is equal to the horizontal tail arm moment. It was ¿

determined as 19.11 in previous stage. It is the distance between the vertical tail aerodynamic centre and wing/fuselage aerodynamic centre. Vertical tail planform area ( S v ¿ is determined by the equation:

S v=

bS V VT 42.4 ×225.1 ×0.08 =¿ 39.95 m 2 = 19.11 lv

Vertical tail aerofoil should satisfy the conditions: generation of the vertical tail lift coefficient ( C L ), ability to generate required lift coefficient with minimum drag coefficient. (Sadraey, 2012). Therefore, the vertical tail aerofoil has been determined as section 6.8.2.4 NACA0012 Vertical tail aspect ratio ( AR v ) has been determined 1.90 through the consideration of advantages and disadvantages. If it is too high, it causes poor lateral aircraft control and prone to fatigue and flutter. However, it could be a good solution to avoid a deep stall because when aircraft stalls, it can keep the horizontal tail out of the wing awake and it is aerodynamically more efficient. (Sadraey, 2012). Vertical tail taper ratio ( λ v ) it is derived by the equation v

2

(1+λ v + λv )÷ (1+ λ v ) MAC= root chord . 4.56m was used as the value of root chord. MAC value was derived 2 × ×¿ 3 by the calculated area and assumed aspect ratio. Vertical tail taper ratio has been derived as 0.97 through this equation.

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Vertical tail incidence angle ( iv ¿ is zero because vertical tail is not required to generate any lift to maintain directional trim in normal flight. This is because the components (engines, wing, horizontal tail and fuselage) that affect directional trim are designed as symmetric about xz plane. (Sadraey, 2012). Vertical tail sweep angle ( Λ v ) is determined as 35 degrees. This is because the jet engine aeroplane that can carry similar maximum take off weight is 35 degrees. Vertical tail dihedral angle ( Г v ¿ does not exist because the aircraft only has one vertical tail. Vertical tail span ( b v ¿ , root chord ( C v ¿ , and tip chord ( CV ¿ , and MAC v can be derived by the root

2 v

bv b AR v = = Cv Sv

tip

Cv

2

1+ λ v + λ v 2 ) , S v =bv C v , λ v= , C v= C v ( equations of Cv 3 1+ λv The results are given as: Vertical tail wing span ( b v ¿ : 8.71m, vertical tail wing chord: 4.58, vertical tail root tip

root

root

chord: 4.56 vertical tail tip chord: 4.42

3. Conclusions Wing design can be summed up is few points. Some parameters are selected from a range of numbers and some are designed from scratch. The design process is all about 18 parameters. Following the factors, between choosing from already available data what will best fit the design requirements or designing the component from scratch. Designing the most important aircraft component is not an easy task to get along with, taking care of each detail is the key to come out with the best results as it will have an impact on other aircraft components.

4. Recommendations Further improvements could be made by researching more aerofoils and more about trim conditions, this would have allowed more suitable and optimised settings for the designed aircraft as well as more analysis about stall recovery and exact t...