2008 SAE Aero Design: Cargo Plane Preliminary Design Review

Transcription

1 2008 SAE Aero Design: Cargo Plane Preliminary Design Review Written by: Jeff Gibson (Team Leader) Jennifer Allison Dan Denmark Ray Klingerman Kathleen Murray Steven Tucker Joe Walk Submitted to: Dr. Sepri Date Submitted: October 26, 2007 Course Titles: Senior Design: MAE

2 Table of Contents Section Title Page Number Need... 2 Problem Statement and Objectives... 2 Goals... 2 Evaluation Criteria... 3 Information Search... 3 Background... 3 Aerodynamics... 4 Wing Structure and Construction... 8 Fuselage and Landing Gear Control Systems Economic Analysis References Appendix A: Schedule Appendix B: Multi-Disciplinary Teams Appendix C: Life-Long Learning: Appendix D: Matlab code for take-off analysis: Appendix E: Selig 1223 CFD Analysis:

3 Need There are two major needs to be addressed by this project. The first need is to compete and do well at the competition. Because Florida Tech has a sparse history in competing in the SAE Aero Design competition, one of the purposes of the project will be to gain more recognition for the school through participation in the event. The second need is the indirect need to maximize carrying capacity for aircraft. This project will act as an exercise in optimizing certain design aspects toward this goal. Many of the concepts and techniques used to design the aircraft can be applied to other real world design scenarios, such as military cargo planes, or passenger airliners. Problem Statement and Objectives This design project has several main objectives that the team will pursue which consist of: Goals 1. Designing and creating an RC aircraft that meets the requirements to compete in the regular class SAE Aero Design 2008 competition. 2. Competing in the SAE Aero Design 2008 competition. The following are the main design requirements imposed by the competition rules and guidelines supplied by SAE [1]. 1. The aircraft must operate using an unmodified OS.61FX engine and E-4010 muffler. 2. The aircraft must be able to take off within a distance of 200 feet and land within a distance of 400 feet at maximum payload. 3. The overall aircraft dimensions (Length + Width + Height) must not exceed 175 inches. 4. The fuselage must be able to fully enclose and support the rectangular cargo box measuring 5x5x10 inches. 5. The aircraft must have a gross weight (including max payload) of no more than 55 pounds. In addition to the design requirements necessary to meet the competition guidelines and rules, the team has defined several self-imposed design requirements: 1. In order to place well in the competition, the final aircraft must be able to achieve a successful flight with at least 35 pounds of payload at sea level density. It will be the main goal of the team to maximize this payload as much as possible while keeping with the design limitations. 2. The aircraft must be able to make a turning radius necessary to make a full circle of the airfield without entering any of the no fly zones. 2

4 3. The aircraft must be easy to control, such that an experienced RC pilot can fly the plane with little difficulty and little practice with the aircraft. 4. The aircraft must be able to sustain the extra load factors of the maneuvers necessary to meet the mission requirements of the competition. 5. The aircraft must be able to sustain the impact of landing and maintain its structural integrity. 6. To ensure reliability through the course of the competition, the aircraft must be easily repaired in the event of a crash or structural failure. 7. To score well on the design category of the competition, the team must develop an equation to predict the maximum payload based on density altitude within 2 pounds. Evaluation Criteria The main criterion for evaluating the success of the project will be the overall placement in the competition. A secondary criterion will be whether or not the team's initial goal for maximum payload is met. Information Search There is a multitude of sources being used throughout the research and design phases of this project. Two major sources provided a great deal of information pertaining to the design. The first is an AIAA journal by E.V. Laitone pertaining to the tandem wing design ([2] Laitone, E.V. Prandtl s biplane theory applied to canard and tandem aircraft AIAA Vol 17, No4, April 1980 pg ). This paper gave the necessary equation for downwash angle, shown in the Aerodynamic Analysis section that allowed the team to develop a model for the aerodynamic forces on the second wing. The second is a white paper published by Leland Nicolai on the subject of aircraft analysis and design as it pertains to the SAE Aero Design competition. ([3] Nicolai, Leland M. Estimating R/C Model Aerodynamics and Performance. June 2002.). Using the information obtained from these two sources, a mathematical model for the takeoff distance was developed, and was used as a method for comparing the performance of tandem wing geometries and configurations. Background The challenge that has been set before the SAE Aero Design competitors is to design and construct a radio control aircraft that can carry a payload successfully. The purpose of this challenge is to offer students an opportunity to apply the knowledge and skills they have been supplied with to a real life situation. Each team entry must follow specific guidelines for the construction of their aircraft in order to qualify for competition. This provides experience with working in a design group atmosphere and with following specific instructions and deadlines. The direct need of the project involves the need for Florida Tech to enter a team into the competition, and place well. There are other secondary needs of the project as well, particularly those which revolve around the military and civil applications of heavy 3

5 lifting aircraft in industry. Competing in the design competition acts as an exercise in creating the most efficient and effective design possible, and may also spur innovation and research into new design concepts. The 2008 Cargo Plane team spoke with the 2007 team to discuss the difficulties that they have faced in their project as well as what their design entailed and their reasoning behind their design decisions. The 2007 cargo plane team faced time constraint difficulties and failed to stick to their schedule. The team s construction is late and some parts are still not finished. This is due to the fact that the designing portion was finished past the scheduled time. The 2007 cargo plane also discussed their design for the plane. An aspect ratio of 15 has been chosen as the best compromise of various factors. The 2008 team will reanalyze this design based on the changed 2008 requirements in order to obtain maximum performance. For practical reasons, the center of gravity of the entire plane should be vertically aligned with the aerodynamic center or with the pressure center of the main wing. The nose then has to be far enough from the center of gravity; this way, the weight which would have to be added in order to compensate the weight of the tail. The size of the current senior team s plane is constrained by the elements which have to be placed inside the fuselage, such as the cargo bay, battery pack and receiver from the RC unit, fuel tank, 2 servos for the elevator and the rudder. Aerodynamics The wing design is extremely important to how an aircraft performs and a great deal of time must be spent to perfect the design. Everything from the shape of the airfoil, the length to surface area ratio of the wing, the twist in the wing, and even how far back the wings are swept must be taken into account for the design. The decision to choose an airfoil for the main wing would be based on characteristics that were considered most important to the design. The airfoil s Lift-to- Drag ratio, maximum lift and ease of construction. However, the main characteristic examined was the airfoils coefficient of lift (cl) vs. both the coefficient of drag (cd) and the angle of attack (α). Through the comparison of data analyzed using the XFLR5 program, which is based off of the XFOIL program, the s1223 airfoil was found to be the superior airfoil design. The aerodynamic coefficients of the Selig 1223 airfoil, as modeled in XFLR5, can be found in Appendix E. This data was used during the analysis and design of the wings, and will be used to precisely model the aerodynamic performance during future design. The Selig 1223 airfoil, which is shown in Figure 1, has been chosen by the team as the airfoil that is to be used. The team compared several different airfoils typically used at the competition; two of these are the Selig 1223 airfoil and the FX airfoil. A plot comparing the coefficients of lift at a Reynolds number of 300,000 of the Selig 1223 airfoil to the FX airfoil, shown in Figure 2 clearly shows that the Selig 1223 has a higher section lift coefficient, as well as better stall performance. 4

6 Section lift coefficient Figure 1: Selig 1223 airfoil [4] Comparison of c l at Re = 300, Selig 1223 FX Angle of Attack (degrees) Figure 2: Comparison of the Selig 1223 and FX airfoils As part of the rule changes for the 2008 regular class competition, there is no longer a restriction on wing planform area. Instead, the sum of the overall aircraft dimensions (Length + Width + Height) is limited to 175 inches. Because of this change in requirements, the initial design concept has been altered to a tandem wing aircraft. A tandem wing aircraft is one which has a second lifting surface for added lift and stability. This is similar to a canard design, however the front surface is meant to significantly add to the lift of the aircraft. Figure 3 below shows the Rutan Quickie, an example of a tandem wing aircraft. 5

7 Figure 3: Rutan Quickie [5] By adding a tandem wing, the aircraft will have greater lifting capacity, while not significantly increasing the overall dimensions. The team plans to have the front wing mounted under the fuselage, and the rear wing mounted on top, as far back as possible. This will minimize the effects of downwash from the front wing, and allow the rear wing to produce more lift. Aerodynamic Analysis In order to determine the best dimensions for the aircraft, we designed a program that optimizes the chord length and wing area based on the take-off performance. In particular, the program analyzes the effect the forward wing has on the aft wing. The forward wing causes downwash on the aft wing, which reduces the aft wing s lift and increases its drag. The following equation 1, from E. V. Laitone s paper [6] describes the downwash angle (w) on the aft wing. 1 2y / b y / b1 2g b1 V CL 1 2y / b 1 1 w AR y / b1 (2g / b1 ) / (1) The longitudinal distance between the trailing edge of the forward wing and the leading edge of the aft wing is denoted as y in the above equation. The height difference between the tandem wings is given by g in the equation and b is the wingspan. The subscript 1 denotes the forward wing. J.H. Crowe, in his paper titled Tandem-Wing Aeroplanes, states that the lift to drag ratio on a tandem-wing aircraft is greater than that of a single wing plane. Further, he concludes that the maximum lift to drag ratio will occur when the tandem wings are furthest apart. This is due to the fact that the downwash on the aft wing is lessened the 6

8 further the wings are from each other [7]. Due to this analysis, we have decided that the forward wing will be as close to the nose of the aircraft as possible, and the aft wing will be placed as far to the rear of the craft as possible. To determine the optimal wingspan and chord length, we analyzed how the take-off distance varied with an increased wingspan and increased chord length. The take-off distance is given by the following equation 2, according to Leland Nicolai [8]: S G 2 VTO (2) a mean 1 where VTO 2W / 0.8 S Cl 2 and a g W T D FC W L (2a) / (2b) The mean acceleration is taken at 70% of the take-off velocity. This affects the total lift and drag within the equation, given by the following two equations. 1 L 2 V TO SC l (3) 2 1 and D 2 V TO SCd (4) 2 The coefficients of lift and drag are determined by the aerodynamic properties of the Selig 1223 airfoil we chose. The program solves for the induced angle of attack, which is subtracted from the angle of attack of the wing. This new effective angle is then used to determine the actual coefficient of lift for a finite wing. The coefficient of lift on the aft wing is solved in the same way, except that the downwash angle must also be subtracted from the angle of attack. To determine the optimal chord length of the wing, the program places the leading edge of the forward wing and the trailing edge of the aft wing at a fixed position predetermined by us. These positions are based off the total length of the fuselage. The program assumes that both wings are rectangular and identical in size. Each chord length is increased linearly at the same rate. Because the leading edge of the forward wing is fixed, as the chord length grow the trailing edge moves closer to the aft wing. This occurs in the other direction for the aft wing, where the leading edge moves closer to the forward wing. As the chord length gets larger, you reach a point where the downwash on the aft wing causes the lift to decrease more than the increased chord length causes the lift to increase. For each wingspan, there is an optimal chord length that maximizes the lift and minimizes the take-off distance. The following graph shows how the effect on the takeoff distance as the chord length grows. The data was taken at a wingspan of 8ft and a weight of 45 pounds. 7

9 Figure (4): Take off distance vs. chord length for 8 foot wingspan The optimal chord length changes as the wingspan changes as well. We iterated our program to find the optimal chord length at a range of wingspans between 8 and 10 feet and found that take-off distance decreased as the wingspan increased. The chord length also increased as the wingspan increased. We have decided on a nine foot wingspan due to structural constraints and machining difficulties as the wingspan is increased even higher. At a wingspan of 8.5 feet, the optimal chord length is about 1.21 feet or about 14.5 inches. Wing Structure and Construction The structure of the wings is an area where significant weight can be taken off by choosing the correct material, but the manufacturing methods also have to be weighed into the material decision. The first material that was considered was a foam which would have had carbon fiber skin and a carbon fiber spar. This material had the lower weight that the team was looking for, but it is also a very complicated manufacturing process. The second material that was considered was balsa wood for the ribs, prefabricated carbon fiber tube for the spar, and plastic sheeting called MonoKote for the skin. This method is also light weight, but the manufacturing for it will be much simpler. 8

10 The team has made the decision to use the second set of materials because of the manufacturability. Ribs Manufacturing the ribs for the wing structure out of balsa wood will be done using a method that has been followed at this institute many times before. The ribs will be constructed out of pieces of balsa wood cut to the chosen airfoil shape. The shapes would be cut out by tracing the shape off of a piece of paper using a pen, and then could be cut out of the main sheet of wood using a simple razor blade. A pen (not a ball point pen though) has to be used when tracing the shape because the ink will be able to pass through the paper, and a pencil will leave impressions in the balsa wood that could cause weaknesses in the structural integrity of the part. After the frame is assembled it would be covered with MonoKote by simply tacking the sheet down and then applying heat to it with an iron. The MonoKote provides several advantages over the composite lay up that was discussed previously. First, the MonoKote has a simpler manufacturing process then the composite lay up. Secondly, the composite material is very hard to repair if there is a problem with it, but with the MonoKote all that has to be done is reapplying the heat. If there are wrinkles it is just apply heat, for a hole there is pressure activated MonoKote to apply in the field, and then a patch can be made of the original MonoKote which is then applied with heat. This material choice was made based on the ease of manufacturing, and it allows for varying the spacing between the ribs as the team sees fit. This spacing between the ribs can be determined by several methods including but not limited to the stress that the ribs will experience and the tension needed from the skin to hold it taunt. After a literature search it was determined that the first failure that the ribs will cause is the buckling of the skin as shown below in figure 4. Skin Ribs Figure 4: Primary failure mode caused by rib spacing This failure will occur before the spacing becomes large enough that the ribs will individually support enough force to cause failure. Presently, the rib spacing has been estimated to be between one and a half to two inches, but once we have a sample of the 9

11 plastic sheeting a simple test can be preformed to determine if the spacing can be larger resulting in a lower weight. Spar The design of the spar started with determining an estimation of the loading and size in conjunction with compiling a list of desired behavior characteristics. Estimates for the size and loading came from the constraints placed on us by the SAE competition guidelines for the regular class competition. Those being, that the maximum weight of the loaded plane, including fuel and payload, would not exceed 55 lbs. From this a simple force balance for the plane during cruise will show that the lift on the wings will be 55 lbs. We increased this estimate to 60 lbs to account for the increased lift needed for take off to accelerate the plane up to cruise altitude. This is a rough estimate that has been recently modified but as of yet hasn t been taken into account for the spar analysis. Then we estimated the maximum wing span available to us using the new dimensional constraints given by the competition guidelines. The rules now require that each plane not exceed a 175 inch sum of the height, width and length of the plane, excluding the prop-length. From this a minimum height was estimated based on clearance for our already existing propeller length, giving an approximate height of 18 inches. This left a maximum wingspan of 9 feet, and gave our maximum length for the spar design constraints. Next we compiled a list of goals we felt it necessary to achieve the best performance of the wings. The list consists of the following goals for the spar behavior: Maximize strength Maximize rigidity Minimize weight Maximizing strength is necessary to withstand the most lift force in order to lift the most weight, which is the entire point of the competition. This strength constraint is mainly addressed by material selection but certain configurations also affected this goal and will be discussed later. Because of the long wingspan in order to decrease any vibrations and flapping of the wings, which decreases the lift efficiency, we realized that the spar design would have to maximize the rigidity of the spar geometry and material composition. Additionally the overall design factor of safety 1.2 had to be considered and to increase the conservative nature of our estimations it was decided that we would model the spar based on the assumption that all the lift would be produced by a single wing even though we had previously decided on a tandem wing configuration. Also, the lift would absorbed by a single spar even though a secondary spar will be necessary to handle the torque experienced by the wing and servo/control surface placement. Using these constraints several geometries were discussed as possible candidates for the spar geometry. Such discussion included the idea for an I-beam configuration manufactured using carbon fiber strands, which was rejected because we felt our fabrication experience was to low to produce something reliable within the time need to complete construction. Also, the I-beam configuration is generally used when the 10

12 bending direction is fixed to maximize the I-beam properties, but our wing will be bending due to drag and lift and the resultant bending plane will shift as velocity is varied and the I-beams behavior will be difficult to predict. Another possibility considered was a square wooden beam with carbon fiber re-enforcement, which was rejected because the square would create stress concentrations at the corners and wouldn t minimize weight. Also, the square wouldn t be a perfect square and would there fore have a preferential bending direction where it would be less resistant to bending and such a direction may not be immediately known and could result in a weak wing. Finally, a tube/cylinder design was decided upon for several reasons. First, the tube maximizes the moment of inertia (I) of the spar while allowing less weight, which both increases strength, rigidity and reduces weight. Secondly, a tubes bending characteristics, namely the moment of inertia (I) is axially uniform, which is great because of the previously mentioned changing directions of the overall bending moment. The cylinder geometry exhibits similar uniform behavior but doesn t minimize weight, however some materials, like wood, can t be made into a 9 foot tube. Then several configurations were considered for how the spar would be attached to the fuselage. Two main configurations were considered; rigidly fixed, and a pinnedsimply supported. It became immediately apparent that the rigidly fixed configuration produced more stress but less displacement and the pinned-simply supported type of configuration was rejected. Below in Figure 5 the final configuration model can be seen. Figure 5: Cantilever beam with uniform load [9] The uniform load was used as an approximation of the lift distribution on the wing even though the actual lift will be more parabolic and be more concentrated towards the half-spar s center. The stress and vertical displacement behavior for this configuration are: 11

13 σ max = (0.5ωl 2 c)/i y max = - (ωl 4 )/8EI [9] ω = distributed load of the lift approx lbs/in (30 lbs/54 inches) l = length of the half-spar (54 in) I = moment of inertia ((π/64)*(do 4 -Di 4 )) E = Modulus of elasticity (Young s modulus) Using this simple model the next stage was deciding the material selection which we selected based on our goal criteria; strong, rigid, and light weight. For a comparison we decided to compare the performance of three different materials; Carbon Fiber, Aluminum 7075-T6, and Ponderosa Pine wood. The Carbon Fiber and Aluminum can both be found in tube configurations and because we want to minimize weaknesses due to poor construction techniques it was decided that the Carbon Fiber would need to be purchased. Therefore it was necessary to run our calculations using cross-sections that are commercially available. Based on the 9 foot wing span and the initial cord-length it was determined that the outer diameter of the tube would have to be around inches. A vender was found that supplied Carbon Fiber tubing with an outer diameter and inner diameter of x inches and with published material properties. This configuration was modeled for both Aluminum and Carbon Fiber as a comparison. The wood couldn t be made into a tube so the cylinder configuration was selected and an outer diameter was determined based on the loading constraints, materials properties and design factor of safety. The results of the overall comparison are as follows: Table 1 Half-spar Material Comparison Results Material Properties Carbon Fiber [10] Aluminum 7075-T6 [11] Ponderosa Pine [12] Yield Strength [kpsi] Allowable Stress [kpsi] Young's Modulus [Mpsi] X10-3 Density [lbs/in 3 ] Estimated cost $400 $60 $50 Half-spar Analysis Results Stress [psi] Maximum deflection [in] Weight [lbs] All of the above materials would be capable of bearing the anticipated loading conditions assumed for the comparison, but the wood weighs the most and the diameter needed to sustain the structure was found to be 1.16 inches; greater than our in outer diameter constraint. The Aluminum is also heavier than the Carbon Fiber and experienced almost twice the deflection. Following up the analysis we generated a comparison matrix as seen below. 12

14 Strength Rigidity Weight Cost Availability Total Carbon Fiber Aluminum 7075-T Ponderosa Pine : 5 Best The selection criteria for the material selection matrix was based on our goals of maximizing strength, rigidity, minimizing weight and additionally the cost and our ability to obtain the material commercially (Availability). The end result was that the Carbon Fiber, even though most expensive, performed the best overall. In order to remain flexible in our design and in case our budget doesn t allow for the cost of the Carbon Fiber our potential fall back options are: A wooden cylinder core with carbon fiber re-enforcement An internally tensioned bow spar made from Aluminum Aluminum tubing with a Blue foam core The problems associated with the wood core and carbon fiber re-enforcement would be modeling the behavior under loading. How the carbon fiber would affect the woods strength and rigidity would be difficult to estimate and testing would be our best option to determine the performance of such a combination. Also, manufacturing would be more difficult and take more time in addition to the time taken up by the testing. ANSYS analysis may be possible but there wouldn t be any comparison for the results to determine if they were valid. The use of Aluminum with a Blue foam core would have the same complications. The internal bow spar design is a concept that could increase the performance of either the Carbon Fiber or the Aluminum. The main concept behind the bow spar is the presence of a wire running through the hollow tube core diagonally from the bottom center the wing to the upper top of the wing tips. The wire will hopefully diminish the deflection of the wings tip under loading and it is anticipated that it may change the bending behavior from that of a cantilever beam with one fixed end to that of a cantilever beam with one fixed end and one simply supported end. This should produce a bending in the middle of the spar that will considerable less than the deflection experienced with out the wire. Figure 6, seen below, is a rough diagram of a half-spar using the bow spar concept. 13

15 o θ T Lift Force Figure 6: Bow spar concept diagram The wire in bow spar (seen here as the red line) will resist the deflection in the tip caused by the lift on the wing. However, the angle theta (θ) is very small (about 0.53 o for our current geometry) and results in high levels of tension in the wire. In order to get a maximum level of possible tension in the wire we assumed that the half-spar was totally rigid and pinned at the connection to the fuselage, transferring all the lift force to the wires tension. From a simple static analysis of the spar the maximum possible tension was found to be: ΣMo = 0 = Lift*27 in T*sinθ*54 in o T = 1620 lbs This is a considerable amount of tension and would require an especially strong cable. Fortunately in our research we found a vender (Fibraplex Corp.) that sells Carbon Fiber Tow; a string made from carbon fiber that has a 750 lbs [13] breaking point and very low weight for $0.40 per foot. Three sets of this cable would allow for over 2100 lbs of tension in the bow spar. The use of this material is expected to allow no displacement of the end point of the spar and can hopefully be modeled using the fixedsimply supported cantilever beam with uniform loading model. Also, it may be possible to analyze the behavior of the spar using ANSYS in comparison to the cantilevered beam model. In the end the best thing to do will be to test the spar and determine if its behavior is the best fit for our plane. If the half-spar does behave the way we expect the model will be similar to that of Figure 7, seen below, and the maximum displacement will become approximately inches in comparison to inches. 14

16 y max = (ωx 2 /48EI)*(l-x)*(2x-3l) Figure 7: Bow spar approximate behavior fixed-simply supported Cantilever beam [9] What s not taken into account by this modeling is the axial component of the tension acting on the spar. Realistically the tension in the Carbon Fiber Tow won t even approach the 1620 lbs maximum tension but will be much lower and will hopefully not greatly affect the integrity of the design. In order to determine if it will have a great impact the static analysis needs to be redone with the beam fixed and then solved using the combined displacement of the wire and beam to determine the overall behavior. Then next phase of design will be the tensioning mechanism and the attachment to the fuselage. Initially the tensioning mechanism seems like it would need to be able to with stand a very large tension in the string, but our intention is to connect the Carbon Fiber Tow to both ends of the wing so only a single tension will run through the entire string and this will be slightly tensioned at the center of the wing. For a quick mental comparison take a string and place it under high tension; it s then possible to increase the tension without much effort by placing your finger on it and slightly displacing it. Essentially that will be the same thing being done by the tensioning mechanism. The displacement will be almost perpendicular to the tension so only a small component of the tension will be absorbed by the tensioning mechanism. Also, initially we won t tension the string beyond removing slack out of the line to help inhibit vibration and jerking of the wing. Further schematics of how the whole spar will look like and operate will created in the next phase of the design. Fuselage and Landing Gear The fuselage provides the majority of the aircraft s structure and integrity. The fuselage also houses all the control systems, the engine, and the cargo bay. Due to the weight of the items within the fuselage, as well as the weights due to the wings and tail, it is necessary for the fuselage to be constructed out of a durable material. The location of the center of gravity is therefore estimated to be located within the cargo bay. The exact location shall be determined once the mass of the components held within the fuselage are known or can at least be approximated based on comparison to similar components. 15

17 In accordance with the requirements from SAE, the cargo box being employed for the plane will be an enclosed rectangular box measuring 5x5x10, the minimum allowed parameters for the box. During the competition, the box will be measured and tested to prove that it can be removed and re-inserted easily. This is done to show that the cargo box does not add any strength or integrity to the structure of the airframe, but that it can be secured to the airframe to prevent the box from falling out during flight. Another requirement is that the box supports the weights placed inside it as a homogeneous mass. This can be done either by having the weights wedged in the box or having holes cut into the weights and securing them with posts to the box structure. The design that is being considered is a box with two interlocking pieces made of either balsa or basswood. Each piece in a simple half box form that will slip together and be locked into place by aluminum posts that screw into the bottom half and stick out the top half, where they will be secured with wing nuts. The basic designs for the top and bottom half can be seen in figure 8 below with the approximate locations of the post holes. Figure 8: Model of the top and bottom half of the cargo box designed in Pro-E by Jennifer Allison The choice of material for the cargo box is still to be finalized, but the thickness of the material has been determined to be 3/16. Taking this into consideration, the wood used for the cargo box must be durable enough to be able to support approximately 25 pounds while having a 3/16 thickness and being light enough for flight. Although the balsa wood is much lighter than the basswood, the basswood is a more logical choice based on general durability and strength, despite being thin. For example, the compressive strength of balsa is MPa while the compressive strength of basswood is between MPa parallel to grain. The flexural modulus should be considered, although it is assumed that flexing will be prevented by the support of the airframe against the cargo box. The support posts for the weights are to be aluminum based on machinability and availability. To aid in the anchoring of the posts, an additional basswood block can be added to the base of the box beneath the holes or these blocks can be attached to the airframe only. This second option would require for perfect alignment between the box s 16

18 holes and the locations of the blocks. The full assembly can be seen in Figure 9, without the posts and wing nuts. Payload plates Figure 9: Model of fully assembled cargo box Designed in Pro-E by Jennifer Allison Another key part to this project is the payload plates. These plates are to be the weights carried in the cargo box inside the plane. According to Section in the SAE AeroDesign Rules for 2008, every team must provide their own plates. [1] This allows for some flexibility when designing the box and plates themselves. Our team has decided upon a simple rectangular design that will fit snuggly into the cargo box. They will also have two holes drilled in them to allow for the support pegs to hold them in place. The approximate surface dimensions of the plates, for now, are x The thickness of the plates will vary according to the type of material used and which weight increment each plate is to have. The material type for the plates is going to be metal, but the type of metal has yet to be determined according to availability and price. The different weight increments are four 1lb plates, two 5lb plates, two 10lb, and one 20lb. The reasoning behind the different increments is so that during testing, each flight can have small increments of weight added on. This will assist in testing the fuselage structure s stability and the overall aircraft performance. Landing Force Calculations The force that the plane will experience upon landing is much higher then the force that it will experience while resting on the ground. This dynamic force was the force that was used to model the landing gear in ANSYS, because this should be the highest force that the plane will experience. A Matlab program was created using equation 5 so that the various parameters could be altered easily. 17

19 h F l m (5) x Where F l is the force upon landing, m is the weight of the aircraft, h is the altitude that the aircraft is landing from, and x is the distance that the plane travels on the ground. This equation assumes that the force felt upon landing is absorbed over a distance x rather then all at one point. This will be accomplished by the flex in the landing gear and the damping associated with the rubber of the wheels on the landing gear. Below in figure X is a sample output from the Matlab program that shows how the landing force dissipates the longer the x component is. Figure 10: Theoretical landing force as a function of landing distance This landing force was used in the ANSYS calculations by taking x = 100ft so that we have a conservative number since the allowed landing distance is 400 ft. Landing Gear Three design concepts have been produced for the landing gear. All of the concepts are of a semi-circular design in order to reduce the number of sharp angles. The decrease in number of sharp angles prevents having more areas of concentrated stress. The bowed shape of the landing gear also allows the structure to flex when force is applied as the aircraft lands. All three of the concepts are to be made of aluminum. ANSYS analysis was done in order to determine the displacement and the von Mises stress of the landing gears when an upward force, the force created during the aircraft s landing, is applied. Design concept 1, shown in figure 11 is designed to be attached to the top of the fuselage. This concept was designed in order to provide a 24in wide spacing between the wheels and maintain a constant radius of 12in and is 1in wide. It is 12in high in order to leave clearance between the propeller and the ground. The fuselage is going to be 6in from top to bottom and then the propeller will reach another 4in below that, leaving a clearance of 2in between the ground and propeller tip. Since the wingspan of each of our 18

20 wings is approximately 9ft in length a wide wheel separation is needed in order to keep the plane stable while it is touching the ground. Also the constant radius is necessary to evenly distribute the load that is applied to the landing gear as it contacts the ground. The maximum displacement of this concept is in and its maximum von Mises stress is 11913psi. This concept however has several flaws. One of these flaws is the fact that when attached to the fuselage, the large cross sectional area will cause more drag. It also weighs 2.122lbs which is a significant weight addition that is not needed. Also, placement of a top mounted landing gear is more difficult than a bottom mounted gear. The location of the center of gravity may place the landing gear at the same point as the lower wing, which it cannot pass through. Figure 11: Landing gear concept 1 designed in Pro-E by Raymond Klingerman To remove the problem that the top mounted landing gear provided, a bottom mounted, constant radius alternative was designed. Concept 2, shown in figure 12, is 6in high, in order to leave the necessary 2in of clearance between the ground and the propeller. This concept is much lighter than the previous, weighing at approximately 0.8lbs. The flaw to this concept is the fact that in order to keep a constant radius of 6in the separation between the wheels is only 12in, causing the aircraft to be rather unstable while on the ground. The ANSYS analysis performed on this concept gave a maximum displacement of in and a maximum von Mises stress of 15685psi. This tells us that the this design deforms less than the taller and wider design but the stresses are greater in the landing gear due to less area to disperse the stress through. 19

21 Figure 12: Landing gear concept 2 designed in Pro-E by Raymond Klingerman Concept 3, as shown in figure 13, is a combination of the two previous designs. It is 6in high, 1in thick and has a separation between the wheels of 24in. This allows it to be mounted anywhere under the fuselage but still gives us a wider separation, which allows for better stability of the aircraft while on the ground. This concept weighs the least amount at approximately 0.2lbs. This concept however does not have a constant radius causing it to have higher stresses acting on it as the ANSYS analysis shows. The analysis showed that the concept will experience a maximum von Misses stress of 48196psi, which is much higher than the other two concepts but still under the breaking stress of the aluminum. The displacement of this concept was also greater, at in. This displacement is allowable due to the fact that there is 2 inches of clearance that allows for such bending of the landing gear. This concept even though it has a higher stress and deflects more appears to be a good choice as far as stability and attachment reasons. Also, the larger displacement will allow the shock of the landing to be absorbed by the landing gear instead of dispersed mostly into the aircraft protecting it from possible damage. Figure 13: Landing gear concept 3 designed in Pro-E by Raymond Klingerman 20

22 Our third wheel design has not yet been determined. More analysis of where the center of gravity of the aircraft is located is going to be done in order to determine whether to have a nose wheel or a tail wheel design. At this time it is anticipated that a tail wheel will be used due to the fact that our second wing will be attached directly to the top of the vertical stabilizer, putting a large amount of mass at the back of the aircraft. Control Systems In all remote controlled planes the control systems are very important for the propulsion and maneuverability of the aircraft. The control system of an RC plane is the mechanical and electrical (see Figure 6, below) parts that control the throttle, and the movement of the ailerons, elevators, and rudder. The control systems are a set of electronics and mechanisms including: Remote controller Digital servos Pushrods/linkages Battery packs Wires/switches Figure 14: Electronics for the control system [14] The remote controller will be a typical RC plane four channel controller that has a channel each for all three degrees of freedom (pitch, up/down, left/right) experienced by the airplane and a final channel controlling the throttle. It will be necessary to make sure that the joystick motions for flying are setup in a typical manner to make it familiar for the pilot to operate. The standard setup used in the US is currently ailerons and elevators controlled by the right joystick and throttle and rudder controlled by the left joystick. 21

23 The competition rules require that the controller transmit according to the FCC and Academy of Model Aeronautics 1991 standards and recommend a 2.4 GHz system to avoid interference. The standards consist of a list of frequencies for different channels in the 72 MHz, and GHz bands, while the average controller available for purchase is generally in the 72 MHz band. The controller will determine the rotation of a set of servos controlling each degree of freedom for the aircraft. There will need to be one servo for the throttle, two for the rudder, one for each elevator, and one or two for each aileron depending on the size and expected forces working the ailerons, the total being 7-9 servos depending on aileron size. Servos are small electric motors than can be digitally programmed to have an output rotation of a specific degree and are controlled by the joysticks on the remote controller. This means that if an object to be moved, such as a rudder, which has no need to fully rotate, it is not necessary to have a motor that rotates a complete 360 degrees. If a motor that rotates 360 degrees is used, a complicated four bar crank rocker mechanism would have to be designed in order to get a small output motion. The servos that are going to be used can be programmed to rotate a certain degree setting to provide the small output movement of the rudder. The servos are programmed by a handheld servo programmer that allows the user to easily and digitally set the specific rotation angles of the servos. The servos will most likely be purchased from or donated by a local hobby shop. The servos available for purchase have a torque output range of oz-in for operation using a six volt battery. The respective weights for the servos are respectively oz and are of varying dimensions and prices. The placement of the servos and which size will depend on how much torque needs to be used and for how long that torque needs to be applied. For instance the rudder will be operating a lot of the time during flight and as the power drains from the batteries the torque output will drop considerably. It will probably be necessary to test the actual torque output and its change over time using a setup similar to that seen below in Figure 15. Servo Scale Figure 15: Servo torque test setup [15] Servo placement will consist of one servo connected to the engine s throttle, which will open and close the intake valve to increase or decrease the power output of the engine. The servo for the throttle will be located inside the fuselage directly behind the motor. The servos for the elevators will be located on the underside of each of the horizontal sections of the tail and will be connected to the elevators using pushrods and linkages. The two servos for the plane s rudder will be located low, on both sides of the 22

24 vertical section of the tail. The rudder needs two servos in order to allow it to cover the full area of rotation necessary for the planes handling. Two or more additional servos will be necessary for ailerons and will have to be located within the wing s cross-section, most likely on a fortified rib. Pushrods are metal dowels, of a small diameter, that are used to convert the rotation of the servo to the output motion needed for the object it is attached to. The pushrods that are going to be used are going to be made using a small diameter aluminum dowel that well be cut to the necessary size. The pushrods are connected to the servos and the flaps by linkages. The pushrods are bent at the end where they connect to the servo and are placed in a hole in the linkage that is glued to the output shaft of the servo. The other end of the pushrod is glued into the linkage that is connected to the flap. The linkages that are going to be used on the plane are going to be prefabricated, store purchased, or donated, typical RC plane linkages. They are typically made of plastic and are connected to the plane by screws, adhesive or a combination of the two. The connection of the flap, by the use of linkages and pushrods, to the servo are shown in Figure 8. Figure 16: Connection of a flap to a servo using linkages and a pushrod [16] The battery packs and wiring are going to be store bought or donated and have to be at least 500 ma-h, as stipulated by the SAE requirements. For the initial design it is expected that two 6 volt, NiMH battery packs will be used to power the five servos. The battery packs will be located in the fuselage and will be connected to the servos by the wiring harness. It is also expected that a third battery pack of a different voltage will be used in the remote controller. Design considerations that will need to be taken into account to prevent failure will be the possible overload of the batteries when multiple servos are operated together. For instance when turning the rudder and ailerons will be used simultaneously, and draw a load for four to six servos at the same time. If the batteries are over loaded the servos won t operate correctly and loss of control could occur. Also, the placement will have to be precise in order to have balance and minimize vibrations. Because there is also a slop test performed by the judges prior to each flight to make sure that there is no give in the controls the linkages will have to be tight and properly put together. There is also a 23

25 requirement for servo testing in the design report that will be turned into SAE prior to competition. Note: Some of the items discussed in this section, as stated, are going to be prefabricated, store purchased, or donated items in order to save the team time and money. Economic Analysis Current Budget: Registration fee $ Control systems/electronics $ Carbon Fiber Spars $ Balsa/Bass wood $ MonoKote Film $ Miscellaneous supplies $ Travel cost $ TOTAL $2, Sources of funding: Because this project is being fully funded by the university, pursuit of financial sponsors has been put on hold. 24

26 References [1] Society of Automotive Engineers. SAE Aero Design 2007 Rules and Guidelines. < [2] Laitone, E.V. Prandtl s biplane theory applied to canard and tandem aircraft AIAA Vol 17, No4, April 1980 pg [3] Nicolai, Leland M. Estimating R/C Model Aerodynamics and Performance. June [4] Pisano, Jessica. Heavy Lift Cargo Plane. Stevens Institute of Technology < /graphs.html>. 23 April, 2007 [5] Wikimedia Commons. Image: Rutan Quickie. < September 10, [6] Laitone, E. V. Prandtl s Biplane Theory Applied to Canard and Tandem Aircraft. Journal of Aircraft. Vol. 17, No. 4. April, [7] Crowe, J. H. Tandem-Wing Aeroplanes: An Examination of the Characteristics of this Type of Wing Arrangement. Aircraft Engineering. October, [8] Nicolai, Leland M. Estimating R/C Model Aerodynamics and Performance. SAE White Paper. June, [9] Budynas, Richard: Mechanical Engineering Design, 8th ed., McGraw-Hill Book Company, NY, [10] GraphiteStore.com. < Copyright September, 23rd [11] Beer, Ferdinand: Mechanics of Materials, 4 ed., McGraw-Hill Book Company, NY, [12] Automations, Inc. Copyright September 25 th

27 [13] Fibraplex Corp. < Copyright October, 22nd [14] Fieldman, Jim. "Great Planes Patty Wagstaff's Extra 300S ARF Product Review. Great Planes Homepage. June < September 12, [15] Troy Built Models. "TBM Servo and Servo Extension Testing. < September 12, [16] Aero Protect Corporation. < Copyright April 24,

28 Appendix A: Schedule Table A-1: Schedule 27