Airplane Design

 

The group spent most of the first semester researching and brainstorming the design of the plane in order to maximize the lift of the craft while minimize the weight and the drag. Emphasis was placed on using the most simple design concievable so that the group could easily construct it on a short enough time table. Here is one CAD rendering of the basic design.

Wing Planform Area

The wing is the most important part of the plane since it is the largest source of lift, drag, and weight. The wing planform area is rectangular and straight. A rectangular wing has the advantage of maintaining a constant Reynolds number along the length of the wing. However, a rectangular wing root is generally weaker than a tapered planform due to the proportionally narrower root chord and high bending moments experienced. The center of lift also tends to be further from the centerline, creating more stress across the wing root. A rectangular wing was easier to construct accurately due to the limited resources available to the team. Thus, it was chosen over the tapered wing. The better performance of the tapered wing would have been lost with inaccuracies in its construction.

Horner Plates

The design included the addition of Horner plates to the end of the wings. A regular wing experiences air flow of the high-pressure region under the wing to the low-pressure region above the wing at the wing tips. This flow of pressure hinders the effective lift of the wing tips by decreasing the difference in these pressures. A Horner plate is added to the end of the wing to prevent this type of pressure flow. These plates added a negligible amount of weight and drag resistance to our wings, while greatly improving its performance.

Dihedral Angle

Another improvement that can easily be made to the wings is the addition of a slight dihedral angle. Wings with a dihedral angle have a slight "V" shape when viewed head on. A dihedral on the wings helps to stabilize the airplane for minor rolling motions from side to side. For example, if one side of the wing drops, it decreases the angle of attack with the oncoming wind, which increases the lift, thereby, raising that part of the wing. The other side of the wing is brought less in line to the oncoming wind and drops lower. The dihedral decreases the lift on the higher side and increases it on the lower side and as a result, the plane rights itself naturally. This righting force decreases as the angle off center decreases and goes to zero when the plane is flying straight. The dihedral is proportional to amount of roll that will be self-righted by the wings themselves. As a result of the way that the dihedral works to stabilize the plane, there is a minor loss in the lift of the wings. It is therefore important to have a dihedral that is large enough to steady the plane and yet not so large as to reduce the lift capacity. In the selection of the dihedral angle, research was done on similar-sized model airplanes and it was determined that the optimal dihdreal angle for a high wing is 5 degrees.

Ailerons

The ailerons of a plane are the movable flaps on the trailing edge of the wings that can be raised up or down to help control the plane. They are used both for directional guidance, and can help with stability. Ailerons decrease the lift of the wing in the up position, and increase the lift in the down position. This allows the plane to go into a controlled roll to make directional turns. The National Advisory Committee for Aeronautics (NACA) found through testing, that the maximum rolling moment is almost independent of the size of the chord - as long as the aileron's chord length is over 15% of the wing chord. The aileron chord should be 20% to 30% of the wing chord, with 25% as the recommended percentage. Therefore, our ailerons have a width of 1.5 inches. The length of an aileron should span 40% to 50% of the semi-wing span for normal control. If the ailerons are much longer, they will be liable to bend and deflect at the hinges. Our design thus requires an aileron that is 17.5 inches long due to a semi-wing span of 3 ft.

Wing Materials

There were four main options for wing material. The main criteria for selection were strength, weight, and ease of construction. The first design uses balsa stringers, which involves strips of balsa are glued together into a skeleton frame. The problem with this design is the difficulty of its construction. Although this material is fairly strong and light, if the plane crashes, the wing will most likely shatter. This implies that at least two wings would be needed for the competition. The second option considered was a composite covering a foam core. This was the option that was used by the previous year's group, which resulted in unnecessary extra weight. While a Teflon covering could have been used to prevent the epoxy soaking into the foam, the weight would still be substantial. The third option was a balsa space frame, which consists of a balsa shell outside of a hollow space inside the wing. It would be extremely lightweight and provide an available space to easily place the ailerons and servos. However, it is very difficult to build and would not have the strength that a foam core would have. The final concept, which was chosen, was foam and balsa construction. This is the balsa space frame concept, only with a foam core. The construction is much easier since the foam core can be easily cut using a hotwire and the balsa applied on top of it. It is of lighter weight than foam covered with fiberglass, though it has less strength. However, the strength of the foam and balsa construction is still great enough to endure normal flight conditions. Plus, since construction is easy, it would be possible to have an extra wing for the competition. This design was also proven at the competition last year by many teams, since it provided the best combination of strength and weight.

Different airfoils were evaluated using a CFD program called ModelFoil. Searching its large database of foils, we were able to easily compare their performance. To evaluate the lift of an airfoil at a certain speed, the team took the coefficients of lift, moment, and drag, and entered them into an Excel spreadsheet shown in Table 1. Other variables entered were wing chord and planform adjustment factors.

Table 1: Definition of Variables

Variable	Value	Comments
Speed (mph)	40
Chord (in.)	11.00
K	780	(780 @ sea level, 690 @ 5,000 ft., 610 @ 10,000 ft.)
Span (in.)	72.00	Greatest wing span allowed in competition
S - Wing Area (sq.in.)	792.00	(Estimate) S = Span * Chord
Tip chord (in.)	11.00
Root chord (in.)	11.00
MAC (in.)	11.00	Mean Aerodynamic Chord (MAC)
alpha0 (degrees)	4.00	Section AoA from airfoil plot
Cl	1.44	Lift coefficient at section AoA from airfoil plot
Cd0	0.016	Section profile drag coefficient at Cl chosen from airfoil plot
Cm	-0.38	Pitching moment at 1/4 MAC from airfoil plot
tau	0.17	Planform adjustment factor (from table) - 0.165 @taper ratio = 1
delta	0.050	Planform adjustment factor (from table) - 0.050 @taper ratio = 1
sigma	1.00	1.00 @ sea level, 0.8616 @ 5000 ft., 0.7384 @ 10000 ft.

 

To determine the speed that the airfoil would experience in takeoff/flight, the team reviewed old footage of the competition. By timing the takeoff and estimating the distance, relatively accurate speed estimates were made. Previous competitors were also asked what their average speeds were.

The 11 inch chord length for the wing was a balance between the weight and size of the wing, and the cargo to be carried. At this chord length, the lift generated by the wing is sufficient to lift the weight of the plane and cargo. If the cord length was increased, it was determined that the lift increase would not merit the additional weight added to the plane.

The variables in Table 1 were then used in the formulas in an output spreadsheet displayed in Table 2. The use of this spreadsheet allowed easy changes and the ability to test many different situations.

Table 2: Formulas and Calculations

Result	Value	Formula
Rn - Reynolds Number	343200	Rn = speed (mph)*chord (in.)*K
AR - Aspect Ratio	6.55	AR = [span(in)]2/wing area (in2)
Taper Ratio	1.00	Taper Ratio = Tip chord (in.)/Root chord (in)
Cd	0.12	Cd = Cd0 +[0.318*Cl2)*(1+ d )]/AR
Pitching Moment (oz.-in.)	-1521.07	Pitching Moment = Cm* s *V2*S*C/3519
Pitching Moment (lb-ft)	-7.92
Drag Total (oz.)	43.68	Drag Total = [Cd* s *V2*S]/3519
Drag Total (lbs.)	2.73
Lift (oz.)	517.83	Lift = [Cl* s *V2*S]/3519
Lift (lbs.)	32.36

These numbers represent the ideal operating conditions of the selected airfoil. The aspect ratio shows the relationship between the chord length and the wing span numerically. The taper ratio is 1 because a rectangular planform area is used. The pitching moment is the moment coefficient of the airfoil that will tip the nose downward in steady flight. The tail surfaces will compensate for this moment. The total drag is a result of the drag coefficient that is a combination of the airfoil drag and the induced drag from steady flight. This is the drag caused by the wing alone and does not include additional drag from the body. The calculated lift is the gross lift for the plane and the maximum cargo weight.

The team finally decided to use airfoil CH10 smooth because of its high coefficient of lift and stability for a wide range of Angles of Attack (AoA). Originally, airfoil OAF102 was chosen, but concerns about the durability of such a thin airfoil arose. It was then decided to change plans to use CH10 smooth. This is because it offers nearly the same benefits as OAF102, but with a thicker profile. The design team had to account for the slight increase in drag for the new airfoil.

Fuselage

The fuselage must fulfill several design requirements, the most important of which is to have a cargo area with a continuous volume of at least 300 in3. Another requirement states the cargo area has to be rectangular. These requirements must be met in order for the airplane to compete.

On most other airplanes, the amount of weight that a fuselage holds is another important characteristic that must be taken into account. However, the cargo and components (fuel tank, batteries, RC receiver, etc.) of this airplane do not put any stresses on the fuselage. This is because these components were attached in such a way as to put loads on the two main spars and front aluminum firewall, not the fuselage. The shape, material, and construction method of the fuselage were a balance between the three previously mentioned specifications.

Landing Gear

A tricycle landing gear places the center of gravity in front of the two rear wheels, and the nose wheel is steerable. Having the wheel at the front prevents the plane from tipping forward and ruining the propeller. It also aids in the steering of the plane. However, if the tricycle gear is tipped backward then the tail will rest on the ground with no self-righting. This can be prevented if the position of the landing gear is shifted rearward from the center of gravity by 5% of the mean aerodynamic center.

The tricycle landing gear was chosen due to the stability it provides during take off and landings. The tricycle landing gear was attached to the double spars that provide the strength for the plane. This meant that the landing gear had to be longer since they were attached higher than usual. This produces a high moment at landing, which may result in the splay out of the landing gear. To prevent this, a threaded rod was attached between the wheels to provide support.

Tail Design

Compared to the other design areas for the airplane, the selection of a tail design is very straightforward. Conventionally, a symmetrical airfoil is used in both the vertical and horizontal stabilizers. Since the stabilizers are very similar to the wings, the same construction methods used to build the wings was used to construct the stabilizers; namely the foam core and balsa method discussed in the wing section of this report. The symmetrical airfoil E168 was chosen for this purpose because of its stable characteristics. Another design consideration were the dimensions of the vertical and horizontal stabilizers - which should not be too large in relation to the overall size of the plane. For example, the 2002 Aero Design team made their vertical stabilizer too tall. This created a large control surface and subjected the plane to excessive moments and instability during its failed takeoff attempt in a crosswind.

Here is an isometric view of the final CAD drawing which includes all parts of the finished plane.

This is a zoomed in shot ofthe front firewall. Here one can see the mounting for the engine, fuel tank, and the rack for the reciever and battery.

This is a view which is meant to highlight the double spar body construction. The wings and landing gear are attached to the spars by sandwiching them with contoured plates, and then tightening them down with nuts and bolts.

Here is a view of the tail control surfaces. They are attached to the spars in the same way as the wings and landing gear.

 

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