General Aircraft Design

General Aircraft Design

In this chapter one, can find the general design of the value cargo carrier of the future. In order to determine this general design, mostly the method of Roskam is used (Roskam, 1997). This method is divided into two main design sequences of which the first design sequence is performed. The second design sequence consists of optimization and more detailed design which is out of the scope of this project. The first paragraph describes the internal lay out of the aircraft which differs from the internal lay-out designed in the mid-term phase. Paragraph two, three, four and five describe respectively the following phases of the Roskam method: Determination of general wing layout, sizing of high lift devices, sizing of the empennage section, determination of the landing gear position.

Finally chapter….

Internal lay-out

As mentioned in the introduction of this chapter, the internal lay-out of the airplane has changed during the final design phase. Whereas the initial concept was designed with a 1,70m lower deck and a 3,50m upper deck (similar to the Boeing 747) these decks are switched in the new design. This is done in order to ease the loading of the airplane and also to reduce the loading time. Because the airplane has a high wing, there are no problems obtaining enough ground clearance and therefore the aircraft fuselage can be lower to the ground.

Currently high lifters are used to load the main-deck of for example a boeing-747. Because now the main deck is situated lower the loading of this deck is much easier and will be quicker. More detailed information on the loading procedure and the duration of it is given in chapter XX. Another advantage of the low main-deck is that with a ramp this airplane can be easily converted to a military aircraft which can load vehicles by just driving them into the airplane. More information on the conversion possibilities and the consequences for the design are given in chapter XX.

In Figure 1 a cross-section of the airplane can be seen. This is a cross-section of the front part of the airplane, where it consists of three decks. The main-deck has a maximum height of approximately 3,1m and has a width of approximately 5.5m. This deck is designed to fit standard KLM pallets which are normally stacked up to 3m high and have a width of 2.44m (Air France- KLM Cargo, 2009). Positioning the pallets at the side leaves an aisle of approximately 0.6m which is considered enough for (horse)stewards and pilots to visit and walk through the cargo deck. The upper cargo deck is 1.70m high and is again sized to fit standard KLM (LD9 and LD3) containers. The third deck is used for the cockpit, for the seating of stewards and to include the wing box structure. More detailed information on the wing box size and structural lay-out can be found in chapterXX. The third deck is sized approximately 2.10 m high in order for the pilots and stewards to be able to stand up right. An overview of all dimensions described can be seen in Figure 1. Important to note that the cross-section varies along the length of the aircraft; At the nose gear, the compartment below the lower cargo floor might be a little larger for storing the nose gear. The same holds for the cross-section at the main landing gear. Also after the wing, the third deck stops which causes the cross-section to change. Finally it can be seen that the upper cargo deck is wider than the actual standard containers of KLM. Since this deck also will be very long this makes it possible for KLM to transport very long odd-sized cargo, like for example oil pipes.

Figure 1: Cross-section of the future cargo carrier

Since the cross-sectional lay-out of the aircraft has changed, also the set-up of the pallets and containers changed. Figure 2and Figure 3summarize this lay-out for both the cargo decks.

Figure 2: Container lay-out upper deck (LD9 containers)

Figure 3: Pallet lay-out lower deck (10 ft pallets)

What can be seen from the figures is the maximum number of 19 pallets(10 ft.) and 18 containers(LD9) which can be stored on the aircraft. As in midterm report chapter XX the length on which the pallets can be stored is determined based on the shape of the fuselage. Due to the clearance at the aft section, the lower deck is much shorter than the upper deck. Also the total length of the airplane has not changed during the final design and is still 70m, which will be elaborated on in the following paragraphs on external sizing.

If all of these pallets and containers would be loaded with the maximum amount of cargo they can carry, 6675kg and 6505kg respectively (Air France- KLM Cargo, 2009), the payload would be approximately 240.000 kg. Nevertheless it is impossible to always have maximum loaded cargo containers and pallets. KLM states that most of the time the cargo compartments of their current aircraft are volume limited and not weight limited. Currently KLM uses an average loading density of 166 kg/m3. When applying this density to the total number of containers and their volume, a payload of approximately 110.000 kg is calculated. Therefore these lay-outs of the cargo decks seem to feasible and also offer enough volume for large, low density, cargo

Determination of general wing lay out

In this part of the Roskam method the general geometry and lay out of the wings is determined. This includes determination of some wing characteristics like the sweep angle and the taper ratio but also the positioning and sizing of the control surfaces.

First of all the choice is made for the positioning for the wing and the structural arrangement for that position. As described in the mid-term report this concept uses a high wing because of the operational advantages. The wings will not be strutted so a wing box needs to be incorporated and the upper part of the fuselage needs to be reinforced to carry aerodynamic loads.

From the budget breakdown in chapter XX and statistical data in the book of Roskam (Roskam, 1997) values were found for the wing sweep, the thickness-to-cord ratio and the taper ratio. These values are respectively 35.5 degrees, 0.087 and 0.24. After the determination of the basic geometry of the wing, the lateral control surfaces were positioned and sized. This is again done according to empirical data from the book of Roskam. In this stage the wing contains two ailerons; a high speed and a low speed aileron, and a spoiler. To support the wing structure, also wing spars need to be positioned. Spars are positioned at the front and the rear of the wing and possibly also in the middle.The rear spar has been positioned so it is located just in front of the hinge lines of the ailerons and spoilers. More detail on the structural lay out of the wing can be found in chapter. The final result of all these values can be seen in Figure 4.

The choice for an airfoil is somewhat complicated since new aircraft nowadays use specifically designed airfoils, instead of NACA airfoils. In addition to that, current aircraft often use more than one airfoil over the span of their wings. It is outside the scope of the project to design an airfoil, therefore at this point the choice for an airfoil will not be elaborated on.

The size of the wing is important, not only for creating lift, but also for the storage of fuel. Therefore the fuel capacity of the wing, as depicted in Figure 4,is calculated using an equation presented in Roskam. This equation is based amongst others on the span of the wing and the thickness –to- cord ratio and also accounts for XXXXX.Finally this results in an available wing fuel storage volume of 283 m3, which is more than sufficient if compared to the required volume of 202,5 m3, calculated in the mid-term report.

Figure 4: Layout of main wing

In this stage, the last two wing characteristics decided on are the wing twist and the wing dihedral. Again this is based on statistical data. Since there are no airplanes of this size using a high wing, these values are more of a rough estimate than the others. Finally the wing dihedral is chosen to be -3,5 degrees with an incidence angle of 3 degrees at the root and 0 degrees at the tip.

Sizing of high lift devices

For the sizing of the high lift devices, chapter 7 of Part II of the Roskam series (Roskam, 1997) on aircraft sizing has been used. To size the high lift devices, values for CLmax, CLmaxTO and CLmaxL are used. From the midterm report these are 1.5, 1.9, and 2.3, respectively.

It is first evaluated if the wing planform can achieve the necessary values for CL. Because the horizontal tail plane produces negative lift, the main wing should have a CL higher than the overall CL for the aircraft. Roskam gives a certain factor which can be used to obtain the necessary lift coefficient of the wing. At the same time, a correction has to be made for the wing sweep. When the final CL required is calculated, it turns out that the basic wing design as shown in the Figure 4is able to generate enough lift.

Now the incremental values of the maximum lift coefficient are calculated, in order to obtain the required incremental section lift coefficient. That is, the incremental lift coefficient of the high lift devices alone. When this is known, a flap type, flap angle, and flap dimensions can be chosen. Because of the high incremental lift coefficient needed for such a large and heavy aircraft, Fowler flaps have been selected. They can be placed at the trailing edge of the entire wing, except for the section where the ailerons are located. In take off configuration, the flaps are set at an angle of 20 degrees in order to generate the necessary amount of lift. In landing configuration, the flaps are set to 40 degrees. This does not deliver the necessary amount of lift, so also leading edge high lift devices are required, also known as slats. To obtain the lift coefficient needed, the slats have a length of approximately 10% of the chord. The flap and slat dimensions can be seen from figure X.2.

Sizing of landing gear

Chapter 9 of Part II of the Roskam series (Roskam, 1997) was used to design the landing gear. Because the drag penalty of a fixed landing gear is far too high for the cruise speed of this aircraft, a retractable landing gear will be used. This landing gear will be configured the conventional way, which means a nose wheel is used.

A weight and balance statement (see Appendix X) has been made using Chapter 10 of Roskam (Roskam, 1997). From this the position of the different wheels of the landing gear has been determined; this can be seen in figure X. These positions have been chosen to satisfy the tip over criteria, as mentioned in Roskam. This should reduce the chance of tipping over to a minimum. Off course the aircraft can still tip over to the back if it is loaded incorrectly. Also the ground clearance criteria as mentioned in Roskam have been taken into account. The aircraft should have a clearance of 15 degrees between the aft wheel of the main landing gear and the tail. Also, it should have a clearance of 5 degrees between the outer wheels of the main landing gear and the lowest point of the wing or engine. These minimum clearances and the actual clearances of this aircraft are depicted in figure X.

Because the tip over criteria demand a wheel at approximately 28 and 34.5 meters from the nose of the aircraft (because of the possibility of tipping over to the rear and to the side), a twelve-wheel main landing gear has been chosen. This landing gear has six double wheels in a row, and covers this range since a wheel has a diameter of 50 inch, or 123 cm. This size has been chosen because it is a standard size for large aircraft. It will certainly be able to handle the weight of the plane and because it is already in use, no new tires have to be developed. By using six double wheels the heavy loads of landing are spread out over the structure, and because there is a wheel at either side of the strut, no unnecessary moments are introduced.

The loads to be carried by the landing gear structures can be calculated as follows

(eq 9.1)

(eq 9.2)

The load to be carried by the nose landing gear struts is 461 kN, whereas the load to be carried by the main landing gear struts is 608 kN.

Retracting the landing gears

Still to be done…. (Part IV Roskam)

Stability and control characteristics

In this section the stability of the aircraft is analysed. At this stage the objective is to verify the stability of the aircraft, and not optimize the empennage area for stability. The area of the empennage surfaces will not be adapted in every step due to the short amount of time available, and the fact that the Roskam method is based on statistics and is therefore not very accurate. Important is that this aircraft should have inherent stability, it should be naturally stable, and not need any augmented stability.

Longitudinal stability

First of all, the aerodynamic centre of the aircraft has been calculated, in order to determine the stability margin. The stability margin turns out to be 4.6 meters, which means the aircraft is statically stable.

The next step in the Roskam method is to determine the necessary empennage area in order to obtain a static margin of 5 percent. It can be calculated that the static margin for the empennage area as determined in the preliminary design is 6.5 percent. This means that the empennage area can be reduced. However, this area is not changed because of two reasons; this aircraft has a high wing configuration which reduces the effectiveness of the empennage area slightly, and the Roskam method is primarily based on statistics of low wing aircraft.

Directional stability

The next step in the stability analysis is the directional stability. The value for Cnb is found to be XXXXX, where the minimum overall level of stability is 0.0010 per degree.

To be completed…

Stability in case of engine failure

In case of an engine failure, the thrust delivered by the operating engine will induce a yawing moment on the aircraft. At the same time, the engine that is not operating induces additional drag. This yawing moment has to be corrected by applying a rudder input. Using Roskam’s method, the maximum allowable minimum control speed can be calculated. At this speed, the rudder input needed to correct for the failed engine should be no more than 25 degrees. From calculations it follows that the rudder input required in this case is 19 degrees, which is comfortably within the limits.