AkG Project Report (1)

5/21/2018 AkG Project Report (1)

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PROBLEM

Conceptual design of a regional transport aircraft having following

requirements:

1. Payload Weight 3000 kg

2. Range 2000 km

3. Endurance 5 hrs

4.

5.

6.

Take-off & Landing Distance 1000m7.No. of Passengers = 20

8. Baggage = 15 kg/ passenger

9.No. of Crew Members = 4

10.


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SOLUTION

WEIGHT ESTIMATION

Initial Weight Weight after Warm-up and Take-off Weight after Climbing to cruising altitude Weight after completion of cruise

Weight after loiter

Weight after landingWe consider ATR 42 as our reference airplane and we will try to modify that

design to achieve our goal.

From historical reference we conceive that our design should incorporate aprop-twin engine configuration.

From the specifications of ATR 42 we see that the Aspect Ratio () is 11.1.Hence we choose the aspect ratio of our design to be 12. Moreover the Wetted

area ratio for this type of aircraft is taken as approximately 5( From historical

data), i.e.

Wetted aspect ratio is given by,


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From curves of historical data we can find the maximum lift to drag ratio, i.e.

From historical data we can express various weight fractions during different

stages of flight as

For Range requirements we have,

Range(R) = 2000km = 6561679.8 ft

From BerguetsEquation (For Prop Engines) we have,

We have,

Solving the equation

Considering loiter time to be Hence once again from Breguet equation we derive

After solving we have


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From historical data we already have

Finally it can expressed that

Where Hence 18 of fuel has been utilised during the whole flight and this is a goodestimate. So we will use it for further use.

Now from the equation of dry initial weight estimation we have

Where From historical data we can express the empty weight ratio as

Hence in this case the equation can be written as

Solving this equation we get


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POWERPLANT SELECTION

We will use the power-plants used in the historical reference airplane ATR 42.

The name of the power-plant is Pratt & Whitney Canada PT6.

WING LOADING SELECTION

We will calculate values for wing loading for different conditions during the

flight of the proposed aircraft.

Stall:

We have restriction on the stall speed

We choose stall speed as Hence from the equation of stall speed we have

If we use values of the parameters in the previous equation

Takeoff:

We have the takeoff distance as roughly 1000m or 3280 ft.

Hence takeoff parameter () = 400


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Where, In our case let us consider

In our case we have

As Raymer has suggested we have taken the value as 0.17Hence from the wing loading equation we have Climb:

Assuming climb speed to be 70 m/s we have

We use the equation for wing loading during climb and we get


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[ ] [ ]

After substituting for all the values we have, Cruise:

For the wing loading corresponding to optimal cruise conditions we have,

Using an empirical relationship provided by Raymer we have an estimate of as,

Substituting the value back to sea level values we have

Hence we will use the lowest value

We have the takeoff weight as

= 22660 lb


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Hence,

AIRFOIL SELECTION

We can estimate design lift coefficient as Putting all the values in the previous equation we have

We inspect various airfoil data so that this value lies in the drag bucket. Wechoose the airfoil NACA 64-415.

WING GEOMETRY AND POSITION

We will use a taper ratio of and we have avoided incorporating sweepin our design for the time being.

We have wing area as and aspect ratio Hence

Mean aerodynamic chord is given by,

The wing will be mounted on the lower part of the fuselage, i.e. we will use a

low wing configuration along with a dihedral angle of 6 degrees (As suggested

by Raymer).

We will use winglets at wingtips to reduce induced drag.


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STATIC STABILITY ANALYSIS

The values of various parameters for a typical Reynolds Number of

for the given airfoil is

Hence using the equation for calculating lift slopes for finite wings we derive,

For static stability analysis we assume following values,Static Margin = 0.1 or 10%Volume ratio of the horizontal tail (suggested by Raymer), The initial downwash is given by

Substituting all the values we derive Similarly,

We intend to use NACA 0012 for the design of both vertical and horizontal tail.

Hence we have following data of that airfoil for a As suggested by Raymer we choose horizontal aspect ratio as

AR = 5 and span efficiency factor e = 0.6

Hence for the horizontal tail lift coefficient we have


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Now we evaluate the location of the neutral point using the following equation

Neglecting terms we have,

Hence location of the centre of gravity from the leading edge of the wing is Now we can use these values to calculate other parameters, Where x is the fractional distance between ac and cg

Hence we can calculate tail setting angle Assuming ,

Solving for we have


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TAIL SIZING

We will use a conventional tail configuration using both vertical and horizontal

tail.

For this purpose we will need to find out the fuselage length as suggested by

Raymer.

Hence in our case we get fuselage length L as For our design and

As the engines of our design are mounted in the front we can use tail arm as60% of the fuselage length.

Hence for horizontal tail we have

Similarly for the vertical tail we can write

Finally,

The taper ratios and aspect ratios we intend to use (as suggested by Raymer)

are, Hence the values of chord and span corresponding to these taper and aspect

ratios are

Horizontal tail:


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Vertical tail:

CONTROL SURFACE SIZING

ELEVATOR:

For elevator sizing we would use maximum allowable elevator deflection to be

15 degrees and at this elevator deflection we would like to trim the aircraft

flying with maximum lift coefficient, i.e. and considering noelevator deflection at cruising trim conditions.Hence from curve we can derive that the change in moment coefficientrequired to achieve trim conditions is Where we have considered

as we had assumed static margin to be10%.

Hence we can calculate as But the previous derivative can be written in terms of other parameters as


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Substituting all the values corresponding to horizontal tail we have

From the curve of

(Nelson) we get Hence RUDDER:

We use historical data for rudder sizing and we conclude that 40% of vertical

tail area should be used up for rudder. Consequently


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AILERON:

For sizing of aileron we will use historical data for reference. We will use

aileron from 50% of wing span to 90% of the wing span. The aileron will also

have a chord of 20% of that of the wing.

FUEL TANKS AND FUEL VOLUME NEEDED

For our design we will use Bladder tanks which are made by stuffing thick

rubber bags in the cavities in the wings and fuselage.

We know the density of the aviation fuel is

We will have 18% of our initial weight burnt during the flight as fuel weight.

Hence weight of fuel consumed is Let us consider 10% more fuel for the losses in the tanks.

Corresponding volume is V-n DIAGRAM

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AkG Project Report (1)