DESIGN. BUILD. FLY Design Report. AIAA Foundation. Cessna Aircraft Company. Raytheon Missile Systems

Transcription

1 DESIGN. BUILD. FLY Design Report AIAA Foundation Cessna Aircraft Company Raytheon Missile Systems

2 Table Of Contents 1.0 Executive Summary Management Summary Conceptual Design Mission Requirements Scoring Analysis Design Requirements Concept Generation Concept Selection Preliminary Design Mission Model Design Methodology Aircraft Sizing Aerodynamics Wing Airfoil Selection Drag Prediction Flight Dynamics Center of Gravity and Static Margin Tail size Control Surface Sizing Stability Derivatives Dynamic Stability Propulsion Battery Selection Motor Selection Propeller Selection Structures Other Important Aspects Cargo Bay Drop System Dual Purpose Rudder Servo Detailed Design Aircraft Dimensional Parameters Structural Characteristics Critical Design Elements

3 5.2.2 Load Paths Material Selection Wing Design Fuselage and Cargo Bay Design Tail and Landing Gear Design Critical Load Case Stress Analysis and Optimization Systems and Sub-Systems Selection Aircraft Weight and Balance Flight Performance Parameters Mission Performance Manufacturing Plan Selected Manufacturing Methods Aircraft Manufacturing Process Manufacturing Milestones Testing Plan Drop System Testing Wiffle Ball Drag Determination Wind Tunnel Testing Structural Testing Flight Testing Performance Results Wiffle Ball Drag Testing Systems Functionality Demonstrated Performance References Bibliography

4 Acronyms, Abbreviations, and Symbols DBF Design-Build-Fly C Ltrim Lift Coefficient at Trim WSU Wichita State University X n Location of Neutral Point CG Center of Gravity X AC Location of Aerodynamic Center SM Static Margin C nβ Yawing Moment Coefficient with Sideslip GM Ground Mission n Load Factor M1 Mission 1 V s Stall Speed M2 Mission 2 V c Cruise Speed M3 Mission 3 V lof Lift of Speed RC Radio Control TOW Take-Off Weight Rx Receiver TOFL Take-Off Field Length NiCad Nickel-Cadmium ESC Electronic Speed Controller NiMH Nickel Metal Hydride AWG American Wire Gauge STOL Short Take-Off & Landing MAC Mean Aerodynamic Chord T/W Thrust-to-Weight Ratio WT Wind Tunnel W/S Wing Loading S&C Stability and Control Re Reynolds Number RAC Rated Aircraft Cost C l Lift Coefficient δ e Elevator Deflection C d Drag Coefficient AC Aerodynamic Center C m Moment Coefficient HS Horizontal Stabilizer L/D Lift-to-Drag Ratio C mα Moment Coefficient with Respect to α C Lα Lift Curve Slope C yβ Sideforce Due to Sideslip LG Landing Gear C lr Roll due to Yaw Rate AR Aspect Ratio C nr Yaw due to Yaw Rate α Angle of Attack C yr Side-force Due to Yaw Rate c t Thrust Coefficient C lp Roll with Respect to Roll Rate c p Power Coefficient C yp Side-force Due to Roll Rate RPM Revolutions per Minute C np Yaw Due to Roll Rate J Advance Ratio C lq Lift Coefficient Due to Pitch Rate P a Power Available C mq Moment Coefficient Due to Pitch Rate P r Power Required ALD Actual Landing Distance T a T r mah K v Thrust Available Thrust Required MilliAmp Hours Motor RPMs per Volt Applied 1.0 Executive Summary 4

5 Wichita State University s (WSU) entry into the DBF competition is well poised to take home the 1st place award. The concept is a low wing tractor configuration aircraft with a conventional tail dragger style landing gear. Mission 2 payload is stored internally while Mission 3 payloads are carried externally using a pylon type system. Following an extensive scoring sensitivity analysis, the WSU team decided to focus their design efforts on minimizing empty weight and servo count while maximizing Ground Mission score. This design emphasis yielded an aircraft that is simple, reliable, and has high mission effectivity. Complicated aerodynamic and propulsion system designs were passed over for simple and proven configurations. This was done primarily to allow for concentration on the task of integrating the aircraft design using proven techniques. Emphasis was placed on integrating the structure for weight savings and developing an effective payload loading and release mechanism. The primary factors contributing to the highly effective design is the use of only two servo actuators and one motor controller. The Mission 3 payload drop mechanism is slaved to the lateral-directional control servo. This system, which is detailed later herein, allows the design to operate with one less servo than would normally be necessary. Servo reduction gives a 33% Performance Parameters scoring advantage over another design which uses an additional Empty Weight lbs servo. Other key aspects of the design allow the aircraft to be Number of Servos* 3 loaded in 10 sec which will be key to winning the competition. Ground Loading Time 10 sec Aircraft empty weight has been held to lbs. The aircraft Mission 1 Laps 9 has top speed of 90 ft/sec which will allow for top scores in Mission 2 Time 72 sec Mission 1 and Mission 2. See Table 1 for a brief overview of the Mission 3 Balls 1 aircraft s main performance parameters. Total Mission Score 188 *Number of Servos includes ESC Environmental factors were considered throughout the design process. The WSU aircraft was designed to operate at Timpa Table 1 - Performance Parameters Field s elevation with an ambient air temperature of 100 Fahrenheit with light and variable winds. All performance figures quoted herein are based upon these atmospheric conditions. Because all of the testing will be conducted in Wichita, KS at much higher air densities and winds, performance numbers will be analyzed with atmospheric corrections in mind. At the submission of this report, the aircraft will be undergoing wind tunnel validations. Manufacturing methods and systems evaluations will happen concurrently to ensure that the prototype flying aircraft will require minimal modifications for competition. The story of this year s WSU DBF team has been fraught with disappointing rule changes and subsequent redesigns, but the team is confident that the final product will be worth the effort. The WSU Design Build Fly team is both excited and honored to compete with other young engineers from around the globe. 5

6 2.0 Management Summary The Wichita State University design team is comprised of a core group of 5 senior students, underclassmen and a faculty advisor. Work began on the project in August of 2014 with three separate groups of 5 seniors. Each group produced and presented their scoring analysis, conceptual design, and preliminary design respectively. A final group was selected by the faculty advisor according to a prearranged criterion. This report details the efforts of the group selected to represent Wichita State University in the DBF competition. Group members and responsibilities are shown in Figure 1. A sixth group member was inducted into the team following selection. His efforts were directed at helping to solve stability and control issues with the unconventional rudder only design. The design team outlined a set of rules and a code of conduct before any the beginning of any design work. This allowed the group to proceed under a well-defined set of guidelines. Scoring analysis and conceptual design decisions were undertaken as a group. Advanced and specialized design decisions were accomplished by primary leads. Secondary leads were established to help the primary validate their design choices. All decisions made by the primary are binding and cannot be overturned by the group as along as the secondary concurs with the decision. In this manner, all members of the group were free to work without unnecessary interface by other members. This structure has kept the group functioning smoothly. Pilot Steve Dockery Advisor Dr. L. Scott Miller Aerodynamics Michael Lamb Flight Structures Patrick Clough Flight Dynamics Roy Moye III Jathurshan Manimoliraja Propulsion Aaron Maurer OIA Trey Cleaveland Underclasmen Stirling Duncan Tracy Barry Tai Lam Jacob Sullivan Figure 1 Team Organization Underclassmen are a big part of DBF at Wichita State University. Many of these students sacrifice much of their own time and effort to help make the team successful. Their duties include helping with design 6

7 work, building models, and conducting testing. Conceptual design is another area where underclassmen have helped the team significantly. Their ideas are often untainted by presuppositions, this allows them to create and analyze concepts with a fresh perspective. Based upon previous DBF competitions and lessons learned, the 2015 team jointly agreed upon a project schedule. The schedule consisted of a highly accelerated preliminary and detailed design stage in order to maximize time allocated for flight testing. Delays and contest rule changes, coupled with design uncertainty compounded the risks of an already accelerated schedule. The design schedule quickly became obsolete. Figure 2 shows the expected and actual dates for the design effort. Design Schedule Design Conceptual Design Preliminary Design Detailed Design Manufacturing 3x4 WTT Prototype Final Aircraft Report Entry Form Report Draft Report Editing Report Due Testing Drop Sys Tesing 3x4 WTT Structural Test Flight Tests DBF Competition 8/31 12/6 9/22 1/9 10/15 2/23 9/27 4/5 4/10 4/12 Aug-14 Sep-14 Oct-14 Nov-14 Dec-14 Jan-15 Feb-15 Mar-15 Apr-15 Figure 2 - Design Schedule 3.0 Conceptual Design Conceptual design was undertaken in a three part effort, scoring analysis, concept generation, and concept selection. 7

8 3.1 Mission Requirements Mission requirements are derived via the provided AIAA DBF rules 1. Individual flight and ground missions as well as aircraft parameters are given scoring criteria. Each mission requires the aircraft to perform a specified task which fits into an overall matrix. Critical aircraft requirements explicitly stated in the rules are as follows: Maximum propulsion battery weight of 2.0 lbs. Maximum allowable takeoff distance of 60 ft. Aircraft must carry supplied M2 payload internally Aircraft must carry supplied M3 payload externally Aircraft must drop M3 payloads individually Aircraft must drop M3 payloads over a designated target Ground Mission The Ground Mission is timed loading exercise in which Mission 2 payload is installed and then removed from the aircraft, followed by the loading of Mission 3 payload. An intermission between mission payloads is allowed for the line judge to inspect the aircraft. Ground Mission Score (GS) = Fastest Loading Time (sec) / Loading Time (sec). Mission 1 Ferry Flight The aircraft is required to fly empty, without any payload, as many laps as possible within 4 minutes. The aircraft must also take-off within 60 ft. and land successfully. Mission 1 Score (M1 Score) = 2 * (Number of Laps) / (Max Number of Laps). Mission 2 Sensor Package Transport The aircraft is required to fly 3 laps with a stack of boards 4.5 x5.5 x10 large, ballasted to 5 lbs. The aircraft must also take-off within 60 ft. and land successfully. Mission 2 Score (M2 Score) = 4 * (Fastest Time Flown) / (Time Flown). Mission 3 Sensor Drop The aircraft is required fly as many successful laps as possible. There is no time constraint. A lap is considered successful if one of the supplied wiffle balls is dropped in the designated area (see Figure 8

9 3). Balls are only to be dropped once per lap. The aircraft must also take-off within 60 ft. and land successfully. Mission 3 Score (M3 Score) = 6 * (Number of Laps Flown) / (Max Number of Laps Flown). Non Mission Scoring Attributes Aside from flight and ground missions, the overall scoring is influenced by aircraft empty weight, number of servo actuators used, and written report score. The number of servo actuators is determined by totaling flight control actuators, motor controllers, and other electro mechanical devices. The non-mission related scoring aspects then multiply with the flight and ground scores. The scores are calculated individually and cascade into each other to form the total score. RAC = Empty Weight * Number of Servos Flight Score = M1 Score + M2 Score + M3 Score Mission Score = Flight Score * Ground Score Total Score = Written Report Score * Total Mission Score / RAC 3.2 Scoring Analysis Scoring analysis is used to determine which aspects of the DBF Parameter Value Range provided scoring equation to focus on during design. The analysis Num. of Servos 3 ±1 helps prioritize factors so that intelligent design trades can be Empty Weight 3.0 lb ±1 made. The Kline McKlintock uncertainty method and the Percent Loading Time 55 sec ±45 Change method were conducted by the group. Group members M3 Drops 6 ±5 conducted analysis independent of one another and results were M2 Time 120 sec ±40 then compared for validity. The Kline McKlintock uncertainty M1 Laps 8 ±2 analysis requires the user to establish a range of data for input Table 2 - Parameter Ranges parameters. Input range can have serious implications for the analysis as results can be skewed if the user selects improper ranges. Table 2 shows the selected mean values and ranges used to compare scoring factors to each other. Historical values from previous DBF competitions were selected for this analysis. The values were selected with the knowledge that some teams would design their aircraft for one mission and throw off the scoring weight. Kline McKlintock analysis uses partial derivatives of scoring equation factors to determine their sensitivity. 9

10 Kline Percent Change 4% 6% 26% 1% 31% 32% Servos Weight Loading Time M3 Drops M2 Time 18% 2% 3% 10% 15% 52% Servos Weight Loading Time M3 Drops M2 Time M1 Laps M1 Laps Figure 3 - Kline McKlintock Results Figure 4 - Kline McKlintock Results The Percent Change method used is more linear. Each scoring parameter was stepped by a unit amount and the change to the overall score was averaged for each step over the selected analysis range. Kline McKlintock and percent change methods both produced similar results with only slight disagreement (Figures 3 and 4). Empty weight and servo reduction were both shown to be key design factors followed by ground loading time and missions 3, 2 and 1 respectively. This prioritization allowed the design team to focus their efforts on designing the lightest aircraft with the fastest loading time and a minimal servo count. By using two different methods to validate their scoring analysis, the team was confident proceeding to conceptual design. 3.3 Design Requirements Based upon the results of Scoring Analysis, the group selected design parameters at the maximum ranges. The only factor that was compromised in determining the design goals was number of Mission 3 wiffle balls. Payload loading time was analyzed through testing (described later herein). A trade study between ground loading time and number of wiffle balls determined that the minimum number of wiffle balls would yield the highest score. Selected design goals are listed in Table 3. Design Goals Empty Weight 2.0 lbs Number of Servos 2 Ground Loading Time 10 sec TOFL 50 ft. Vc 90+ ft/sec Wiffle Balls 1 Table 3 - Design Goals 3.4 Concept Generation Each member of the design group generated concepts independent of other members. Concepts were then compared and duplicates were removed. Aircraft configurations were considered as a whole concept. This is a break from pervious design philosophies where each individual component for the 10

11 aircraft is scored and selected separately. The design team felt that the whole aircraft should be modeled as a complete piece. Individual components are often the hallmarks of a certain design and should not necessarily be designed separate from the main aircraft. The three main concepts considered were high wing, low wing and flying wing type designs. Figure 5 shows VSP 2 models of the three final concepts. A Low Wing B Strutted High Wing C Flying Wing Figure 5 - Generated Concepts 3.5 Concept Selection A standardized rubric was developed to rate aspect of the concepts. The weighting of design aspects were chosen to complement those found during scoring analysis. Design Simplicity Ease of manufacture, applicability of basic engineering methods. Multipurpose Structure Reduces structure reduces weight which is a primary scoring factor. Design Flexibility Allows the design team to reconfigure the design in the event of rule changes (which Is historically commonplace) Ease of Loading Allows for fast loading times, which is a primary scoring factor Selection Criteria FOM A B C Design Simplicity 10% Multipurpose Structure 30% Design Flexibility 20% Ease of Loading 40% Score Rank Table 4 - Selection Criteria Table 4 shows the selection criteria used for the final concept. The final concept was selected as the low wing design. Trade studies were conducted to determine propulsion and landing gear configurations. A conventional (tail wheel) design was selected to reduce weight and complexity. Propulsion design was 11

12 tasked with determining the impact of dual engine system vs a single engine. The increased complexity and weight for comparable thrust systems left no doubt that a single engine setup was key. From the beginning of the conceptual design process, it was apparent to the group that an aircraft with more than the absolute minimum amount of servos would not be competitive. A drop mechanism for mission 3 needed to be actuated without adding an extra servo. In addition, flight control servos would be held to the absolute minimum. Only two servos would be used, one for longitudinal (pitch) control, and one for lateral/directional (turning) control. Propulsion control via the electronic speed control (ESC) was mandated in the DBF rules, so there was not any serious contemplation of manipulating that system. 4.0 Preliminary Design The main objective of preliminary design was to size the aircraft and roughly spec components for the final aircraft. Trade studies and an iterative design process is the primary mode of activity. Each of the design leads identified key aircraft parameters, relevant to their aspect of the design, that were required to achieve the performance goals identified in conceptual design. Aerodynamics Group Wing Geometry and Size Influences drag and takeoff performance, constrained to < 60 ft. Airfoil Selection Provides optimal lift and drag polars with various aircraft weights. Aircraft Drag Coefficient Directly influences cruise speed goal of 90+ ft/sec. Flight Dynamics Group Static Margin Effects aircraft controllability and trim drag, cruise speed goal of 90+ ft/sec. Control Surface Sizing Effects aircraft controllability, and mission completion. Dihedral Angle Effects roll/yaw coupling and aircraft controllability. Propulsion Group Battery Size Effects aircraft available and endurance. Effects empty weight goal of 2.0 lbs. Motor Selection Effects aircraft propulsion system efficiency and power output. Propeller Effects aircraft endurance, takeoff performance, and maximum cruise speed of 90+ ft/sec Structures Group Spar Sizing Effect maximum load factors, turn rates and therefore M1, M2, M3 score. Load Distribution Effects primarily empty weight goal of 2.0 lbs. Material Selection Affects both maximum load factors and empty weight goal. General Layout Affects both ease of construction but also Ground Score loading time. 12

13 OIA Group Drop system actuation Directly affects the servo count goal of 2 servos and M3 score. Landing Gear Integration Affects both empty weight and AIAA successful landing criteria. 4.1 Mission Model A basic mission model was developed to outline assumptions and provide for a standard that all design leads would work from. The standard mission was segmented into 5 phases: takeoff, climb, cruise, turns, and landing. All mission phases occur under a set of atmospheric conditions defined by the design team. Standard conditions were derived based upon the outlier condition of a 100 F day at Timpa Field, Tucson, Arizona. Takeoff TOFL was determined using an integration approach. Rolling stock friction was set at.03. V lof was set as 1.2 V s. Aircraft AoA for takeoff is held constant for the calculation since the aircraft tip back angle roughly equals the stall AoA. Winds are only considered for their implications on controllability; credit is not taken for headwinds. Crosswind cases are calculated for 15 ft/sec crosswind component. Climb The climb rates and times are calculated based upon a maximum available power setting. Climb is assumed to peak at 100 ft. AGL. Cruise Level unaccelerated flight is assumed for the cruise condition. All flight mission are to assume maximum power setting. Propellers are selected specifically for each mission and thrust available curves vary accordingly. Turns Turns are assumed to be accomplished at the maximum cruise power setting. Turn rates are calculated at cruise speed and a 60 back angle; approximately 2.0G. Landing Landings are assumed to be power off in case of propulsion system failure. Winds are only considered for their implications on controllability; credit is not taken for headwinds. Crosswind cases are calculated for 15 ft/sec crosswind component. 13

14 4.2 Design Methodology Initial design analysis and sizing was conducted in parallel using the 5 main design areas. Aerodynamics, Propulsion, Structures, Flight Dynamics, and OIA (Other Important Aspects) each conducted design in concert with each other. Methods used during initial design consisted of basic analysis to generate conservative numbers. Once the basic analysis was complete, the group proceeded to a more detailed analysis. The detailed analysis proves more accurate numbers which are then used to refine the design. The following details each segment of the design process. 4.3 Aircraft Sizing An initial empty weight of lbs was estimated for the aircraft by both estimating weight buildup of expected materials and components and utilizing historical data of previous top ranking aircraft in the competition to estimate an achievable payload fraction that would be competative. The aircraft was sized by using Anderson 3 and Raymer 4 simplified performance equations for critical cases and solving T/W as a function of W/S. The critical cases for the mission assigned are takeoff, cruise, sustained turn, stall, and max T/W. The critical design case is mission 2 since it is the mission with the highest payload weight. The aircraft must be able to perform as designed under all circumstances. A maximum T/W of.56 was assumed based off of early battery selection and experimental propulsion data from the Wichita State 3X4 Wind Tunnel as explained in the section 4.6. Stall speed was estimated to be 35 ft/s based off of historical data from the competition and an estimation of a target takeoff speed for a short takeoff given V takeoff = 1.2*V stall. A C LMAX of 1, cruise velocity of 80 ft/sec, and takeoff distance of 50 ft were all also assumed. When these functions are plotted with respect to W/S, the area above the cruise, takeoff, and sustained turn but also below the max T/W and to the left of stall, is the possible design area as seen in figure 6. A design point must be chosen within this are for the aircraft to be able to perform all requirements. The highest performance will come with a design point chosen in the upper right corner of this area. The design point chosen for this aircraft is T/W =.56 and W/S = Given the design point, wing area can be solved for, effectively sizing the aircraft. This gives a wing area of 5.38 ft 2. And aspect ratio of 6.7 was chosen to give a span of 6 ft and a chord of.9 ft. 14

15 T/S T/W vs W/S Take off Cruise 1 Climb Design Area Sustained Turn T/W Max Stall Design Point W/S Figure 6 - Aircraft Sizing Plot 4.4 Aerodynamics The main function of the aerodynamics group is to optimize the aircrafts performance and to predict lift and drag characteristics. The best way to improve aircraft performance is by reducing drag. Drag can be minimized by choosing the correct airfoil, correct aircraft sizing, reducing aircraft wetted area, and overall streamlining the aircraft. Drag is especially important to this competition as discussed in the scoring analysis. In order to have a competitive aircraft, drag must be minimized in all aspects Wing Airfoil Selection Correct airfoil selection is critical for a low drag, high performance aircraft. Keeping flight conditions within a drag bucket of the drag polar keeps parasite and induced drag down, so an airfoil that produces a low C D for the cruise C L is desirable. It is also desireable to chose an airfoil that has a large C lmax, as a high C lmax will shorten the takeoff distance. 15

16 Airfoil Design Requirements The payload weight varries greatly per the given missions. This means that the aircraft will be required to fly at a very different cruise C L for each mission, varrying from to The mission also calls for a short take off distance, so an airfoil with a C LMAX greater than 1 is desiralbe. To ensure performance is represenative to that of the airfoil data, the airfoil must be easy to manufacture. Slowly rounded, thicker airfoils are prefered over thin airfoils with sharp, thin trailing edges. To keep ease of manufacturing high and perfromance hich, an airfoil with a max thickness of 10% to 15% is ideal. Airfoil Selection Process Knowing the design requirements, Seilig s airfoil data 5 was filtered through a in search for airfoils that provide these characteristics. Seilig s airfoil data has a very large number of airfoils that have been experimentally tested at low reynolds numbers, so this data is reliable and very applicable to our scenario. To ease the search for an airfoil that meets all of the design requirements, all airfoils with a C l MAX below 1 were not considered as a large C lmax is needed for short take off. Also any airfoils that had the drag bucket out of the range of the design cruise C L s were not considered. Eventually two airfoils were selected that fulfilled the design requirements and out performed all of the others in these fields. The S8052 and SD7062 airfoils were selected as the final candidates for preliminary design. Each candidate had a tradeoff for performing well in one area (see Figure 7 and Figure 8). The S8052 airfoil has a very large flat drag bucket at a very low C D through the design C Lcruise range, but the C l MAX is only 1.2. The SD7062 airfoil has a C l MAX of 1.6, but the drag bucket curves off through the design C Lcruise range, increasing C D at the lower section of the range. To compensate for the lower C l MAX of the S8052 the wing area will need to increase to maintain design take off performance, but the low drag bucket through all of the cruise C L ranges will reduce the induced drag. The SD7062 wing area will be smaller but induced drag will be higher. A trade study was conducted to compare the drag effects these produce and to see which would overall be better for the given mission. The trade study consisted of modifying the airfoil data and modifying the wing sizing to fit the mission requirements and looking at overall drag produced. The S8052 had lower overall drag, so it is the selected wing airfoil 16

17 2 CL VS α 1.5 C L S8052 SD α (degrees) Figure 7 - Comparison C L vs. α Plot 2 CL VS CD 1.5 C L S8052 SD C D Figure 8 - Comparison C L vs. C D Plot Figure 9 - Two Final Airfoil Candidates: S8052 (Left), SD7062 (Right) 17

18 4.4.2 Drag Prediction Nicolai s drag prediction method 14 was used to predict the aircrafts drag coefficient. A component drag buildup method was also used and compared to Nicolai s results for validation. The values were very comparable. This drag buildup method was done by considering each component of the aircraft and finding its individual drag coefficient and then weighing it with respect to the wing reference area, shown in Figure 10 and Figure 11. The total drag coefficient was found by summing up all of these weighted coefficients. An AVL 7 model was created to validate some aerodynamic values. The main focus of the AVL analysis for the aerodynamics group was the induced drag calculations. The induced drag from AVL was very close to the induced drag calculation from Nicolai s drag prediction. The overall aircraft C Dcruise is.026/.025/.035 for M1/M2/M3 respectively. Component C D,0 % of Total Wing % Fuselage % Stabilizers % Ldg. Gear & Pylons % Total % Stabilizers Landing Gear & Pylons Fuselage Wing Figure 10 - Mission 1 & 2 Drag Buildup Component C D,0 % of Total Wing % Fuselage % Stabilizers % Ldg Gear & Pylons % Wiffle Balls % Total % Wiffle Balls Landing Gear & Pylons Stabilizers Fuselage Wing Figure 11 - Mission 3 Drag Buildup 18

19 4.4.3 Fuselage Aerodynamics To reduce drag produced by the fuselage, a slightly modified NACA 0015 symmetric airfoil cross section was selected. The modifications were made along the bottom of the airfoil to mate with the wing airfoil shape better. The airfoil cross section allows for very smooth flow. A symmetric airfoil was selected to produce minimal lift would be produced at cruise, decreasing excess induced drag. The NACA 0015 was chosen because of the mission 2 payload size limitations. The NACA 0015 is thick enough to allow space for the height of the payload with the chosen fuselage chord length. To further reduce drag, the aft section of the fuselage is tapered, reducing the wetted area. 4.5 Flight Dynamics The stability and controls section of our design group has the responsibility of making sure we have a stable aircraft in flight and to ensure the aircraft can maneuver through each mission with precise control. During each phase of design, multiple calculations were conducted to ensure the correct center of gravity, static margin, tail sizing, control surface sizing, stability derivatives, and dynamic modes were produced by the aircraft Center of Gravity and Static Margin The center of gravity (CG) of an aircraft is a vital piece of information that determines if the aircraft is stable. During the design process, the CG was calculated for each mission using basic moment equations. CATIA aircraft buildups were used to validate the CG hand calculations. Using recommendations from previous Wichita State DBF teams 8 and the Raymer 4 text, a static margin range of 5%-15% was chosen. For all missions, the static margin for the aircraft is within the range of 9%-10% and this range provides adequate longitudinal stability for the aircraft Tail size The horizontal and vertical tails are key components on the aircraft that provide longitudinal and lateral stability. The placement and dimensions for each tail were selected based off of the recommended tail volume coefficients provided in the Raymer 4 text. Each tail also played a role in accomplishing the desired CG and static margin Control Surface Sizing To ensure adequate control of the aircraft, the three main control surfaces need to be sized appropriately. The challenge of the selected concept was the use of only rudder and elevator for the primary control surfaces. During the scoring analysis it was found that the amount of servos used for the entire aircraft deeply impacted the overall score in a negative manner if three or more were used. Due to the mission scoring equation and a desired advantage, a team decision was made to eliminate the use of ailerons on 19

20 the aircraft so that only two servos would be used for the entire aircraft. Since the aircraft has no ailerons, the rolling motion of the aircraft is being created with the combination of an eight degree dihedral of the main wing and the rudder. The dihedral angle of the aircraft wing was calculated using equations from MIT Basic Design Rules 9. The elevator is responsible for the pitch motion of the aircraft and the trimming of the aircraft after it has experienced a pitch disturbance. This control surface was sized with Raymer 4 recommendations of 25% of the horizontal stabilizer chord and 90% of the horizontal stabilizer span. The span of the elevator was ultimately designed to be the full length of the horizontal stabilizer for simplicity and manufacturing ease. Multiple calculations, stemming from Etkin 10, were conducted to ensure the aircraft could be trimmed with the use of the designed elevator and the results are displayed in Table 5 and Figure 12. Also calculations were conducted to determine the elevator deflection required to takeoff for each mission and the results are displayed in Table 6. Mission α trim Cruise δ etrim Cruise α trim Stall δ etrim Stall (deg) (deg) (deg) (deg) Mission Mission Mission 3 (Ball 1) Mission 3 (Ball 2) Table 5 Elevator Trim Deflection Elevator Deflection Angle at Takeoff (deg) Mission Mission Mission 3 (Ball 1) Mission 3 (Ball 2) Table 6 Takeoff Elevator Deflection 20

21 Cm Trim Plot δe=0 δe=-2 δe=-4 δe= CL Figure 12 Graphical Trim Analysis The rudder is responsible for the yaw motion of the aircraft and for the aircraft; it controls the releasing mechanism for the wiffle ball drop system. This control surface was also sized with Raymer recommendations of 25% of the vertical stabilizer chord and 90% of the vertical stabilizer span. In mission three of the competition, our airplane is required to carry wiffle balls and release them during a specific segment of flight. Our team decided the aircraft would use the rudder servo to release the wiffle balls at the appropriate time in flight. Due to this decision, it was imperative that the max rudder deflection allowed be calculated to provide a safe range for deflection of the rudder during our wiffle ball release. The calculated stability derivatives were used in equations from the Roskam 4 text and helped determine a maximum required rudder deflection of 20 degrees. According to the Roskam 4 text, the recommended maximum rudder deflection needed during flight should be less than 25 degrees and our aircraft meets that recommendation. This also concludes that there is still 5 degrees of deflection that can be used to release the wiffle balls during flight Stability Derivatives The stability derivatives of an aircraft provide valuable information on the stability of the aircraft and are helpful in calculating multiple parameters that produce other pieces of important information. These derivatives were calculated using equations from the Etkin 10 and Roskam 11 texts and are displayed in Table 7. After these calculations were completed, the stability derivatives were further validated with the use of Athena Vortex Lattice 7 (AVL) program. Another parameter AVL was able to validate was the aircraft s ability to be spirally stable. According to AVL 7, the aircraft is stable if the spiral parameter is greater than 1 and the aircraft produced a spiral parameter of

22 Imanginary Axis Stability Derivatives C mα (Mission1) Cl r C mα (Mission 2) Cn r C mα (Mission 3) Ball Cy r C mα (Mission 3) Ball Cl p C Lα 5.14 Cn p Cl β Cy p Cn β C Lq Cy β C mq Table 7 Aircraft Stability Derivatives Dynamic Stability The dynamic stability of the aircraft was verified with the use of a root locus plotted in AVL 7 that stemmed from the stability derivatives in Table 7 and the aircraft s moment of inertia values. The root locus in Figure 13 displays the different modes and poles of the aircraft. The spiral pole is close to being positive which is an indication that the spiral mode is close to being unstable. The short period poles lie on the real axis and suggest that this mode is unstable. Upon further observation, the other poles located on the root locus are negative and therefore conclude that the aircraft is dynamically stable. Spiral Dutch Roll Short Period Roll Phugoid AVL Root Locus Plot Real Axis 0 Figure 13 Aircraft Root Locus 22

23 4.6 Propulsion The primary goal of propulsion during the PDR stage was to refine performance calculations and specify propulsion components to be used. Performance requirements and primary drivers were calculated during the CDR stage. By refining propulsion system performance numbers, other aspects of the overall design could be more accurately modeled Battery Selection The first component of the propulsion system to be analyzed was the battery. Because of the emergence of lithium batteries, the technology for nickel type batteries has remained stagnant. The only advances in nickel type batteries have been for low amp draw applications like remote controls. This allowed our team to capitalize on previous research at Wichita State University as well as the testing of past WSU DBF teams. Elite 1500 s were selected by this year s team. They were selected for their excellent performance under high load and ability to handle multiple charge/discharge cycles. 8 It should be noted that Elite 1700 s have better performance than the Elite 1500 s but are no longer in commercial production. 14 cells were selected to maximize voltage and help reduce current draw for the given power required. 14 cells provided the maximum voltage allowable by most motors of the size we were considering Motor Selection Motors were chosen based upon output power. Both geared in-runner and out-runner motors were considered. Table 8 shows the motors combinations considered. A twin engine Rimfire.10 concept was considered because of the possible benefits of vectored thrust, but was soon discarded for its high weight and excess complexity. Geared in-runner motors were considered based upon their popularity in recent DBF competitions. However, despite their popularity, no clear advantage could be found with selecting a geared motor over an out-runner equivalent. The geared in-runners were often heavier and supplied a lower shaft RPM to the propeller, which can be detrimental to efficiency. Given the high power requirements of this competition, larger motors became more viable. Larger motors are more often 23

24 available in higher voltages and efficiencies which work to the benefit of this motor trade study. The selected motor is a Rimfire.15. The motor has a very good power to weight ratio and operates at a reasonably high voltage which improves endurance and reduces cell count. Because the motor will not be running at maximum capacity, it operates at cooler and more efficiently. Motor Style Power Output Weight 2x Electrifly Rimfire.10 Twin Outrunner 650 W (325) 5.0 oz (2.5) Electrifly Rimfire.15 Outrunner 500 W 3.6 oz Electrifly Ammo 28 mm Geared Inrunner 518 W 4.6 oz Scorpion SII kv Outrunner 600 W 4.55 oz Table 8 - Motor Candidates Propeller Selection Because each flight mission has different performance requirements, a propeller is custom selected for each mission. Propeller test data in the WSU 3x4 Wind Tunnel was scored in order to find propellers which had power input requirements that could be supplied by our chosen motor as well as output thrust that could outpace our thrust required. Variables considered in selecting the propeller include static thrust, zero thrust speed, and power required. For mission 1 and 3, a low static thrust could be tolerated as the aircraft was taking off at light weights. However, top speed had to be maximized as mission 1 is speed critical. Therefore a high pitch cruise prop was required. The APC 8x8 provides a thrust available that maximizes near the top end of the flight envelope. See figure 14 for mission 1 and 3 Thrust Available vs Thrust Required plots. Mission 2 required a high static thrust since the aircraft would be taking off at the maximum gross weight. Top speed is still a critical factor, as the mission is timed, but not nearly as important as getting airborne! The APC 10x6 was selected for mission 2 because of it s incredible static thrust. Top end speeds with the APC 10x6 are still comparable to that of the 8x8 and speed should suffer only slightly if at all. See figure 14 for the mission 2 thrust available curve. It should be noted by the reader that the APC 10x6 and 8x8 propeller selections both deliver the same top speed given the current aircraft drag. This is where a distinction should be made between the two. The APC 8x8 propeller has more potential for speed gains in the event that aircraft drag decreases. It should also be noted that for longer duration flights like mission 1, the current draw of the APC 8x8 offers much more endurance and therefore safety margin. In summary, both the need to reduce current draw for mission 1 and the hope for performance gains are the two main reasons multiple propellers have been allocated. 24

25 Thrust / Drag lb Thrust Rqd. Mission 1 Thrust Rqd. Mission 2 Thrust Rqd. Mission 3 Thrust Available Mission 1/3 Thrust Available Mission V ft/sec Figure 14 - Thrust Available vs Thrust Required Plot The overall propulsion system designed for the competition year will provide incredible performance. With an overall system weight of lbs, the system will provide over 4 lbs of static thrust at a power output of 425 Watts. Complete system performance numbers are shown in Table 9. In order to provide a better understanding of performance capabilities, a takeoff performance study was conducted. The results of this study, allowed aerodynamics to reduce wing area and further help reduce aircraft weight. Takeoff distance was calculated using Anderson 3, Nicolia 6, and Integrated Thrust methods, Table 10. The maximum allowable design takeoff distance was set at 50 ft. Wing area and other factors were manipulated until this value was reached. This allowed us to fully optimize aircraft performance and take advantage of propulsion system weight already allocated because of other performance aspects. 25

26 Propulsion System Specifications M1 M2 M3 Motor Rimfire.15 Battery 14 Cell Elite 1500 Voltage 15.4 (14.0 after losses) Propeller APC 8x8 APC 10x6 APC 8x8 Max Current amps 16.5 Endurance 5.5 min 3.2 min 5.5 min T / W Static Thrust 1.75 lb 4.0 lb 1.75 lb Table 9 - Propulsion System Specifications TOFL Predictions Mission 1/3 Anderson 27 ft. Nicolai 35 ft. Integrated Thrust 41 ft. Average 36 Mission 2 Anderson 62 Nicolai 68 Integrated Thrust 50 Average 58 Table 10 - Take Off Distance Predictions 4.7 Structures Structural design involves an iterative process in which all aircraft components are sized, weighed, and analyzed to assure conformance to the performance parameters established by the aerodynamics and stability and controls groups. The structural design group is responsible for making sure the aircraft and all of its components will withstand the loads experienced during all critical flight conditions while maintaining the lightest weight possible. The structural design process involves four major steps shown in Figure 15. The step is to identify the purpose of the part, the surrounding structures, the loads that will be applied, and the location and types of reaction forces that it will experience. The next step is to draft a new geometrical representation of the part that will satisfy its intended function. The part is then put through the structural analysis and optimization process which is outlined in section 5.2. Lastly, the part is checked for manufacturability. Once the part has passed through these steps the design is frozen and will not be changed unless absolutely necessary. Identify Purpose, Loads, & Reactions Draft New Geometry Structural Analysis & Optimization NO Freeze Design YES Can It Be Built? Figure 15 Structural Design Process 26

27 4.8 Other Important Aspects The main goal for the OIA area of design was to ensure the planes ability to handle the missions. This included the addition of a cargo bay and a drop system. The cargo bay would need to be incorporated into the plane in such a way that it did not create a significant drag penalty to the fuselage. The drop system would need to be simple and reliable so that it would ensure a positive release every time. Both of the systems would need to be light so that it does not adversely affect the empty weight Cargo Bay From the mission requirements it was determined that the aircraft would need a large cargo area for the fuselage. This cargo area would need to be able to house the 4.5 x 5.5 x 10 wooden pine block. The block will sit in the cargo area with the 4.5 direction being the z-direction. Orienting the block this way will allow for a better fineness ratio of the fuselage. The fuselage was made to be 6 wide to leave 0.5 on either side of the block to make room for the tolerance, structural components, and electronic components. This tight fit will keep the block from sliding around during flight. In order to ensure a consistent static margin throughout all missions, the cargo bay was positioned so that the CG of the mission two payload would be as close as possible to the CG of the empty plane. The distance between these two points is about half an inch. In order to keep the loading time to a minimum, a cargo door would need to be put in at the top of the fuselage. A top loading design would ensure that the block could be loaded without having to manipulate the plane. The door would be equipped with a quick locking system that would allow the loading team to quickly unlock the door and lock it back in place. The lock would also ensure that the door did not get bumped open during flight. The door would swing forward, ensuring that if there were to be a failure of the lock system mid-flight, the airflow would help to keep the door shut. In order to unload the block quickly without having to squeeze into the sides of the fuselage, light handles will be on the sides of the fuselage that loop down under the block. This will allow the person unloading the block to simply grab the handles and pull out the block, ensuring that minimal time is wasted attempting to get a hold on the block Drop System Initially for mission three, a rail system was designed for the wiffle balls to be held. These rails would hold the balls in the flow, letting drag slide the balls backwards as the balls are deployed. This idea was abandoned after a closer look at the scoring analysis. It was determined that decreasing the amount of time to load the mission three payload was more beneficial to the score than having a greater number of balls. Through practice runs with a mockup rail system and a prototype fuselage, almost five seconds of load time was able to be saved by abandoning the rail system with six balls and using a pylon system that 27

28 only has two balls. This allowed for a faster load time with minimal risk of damaging the deployment system while loading. In addition to the loading time score advantage, the pylon design also offered weight savings when compared to the original rail design. After a mock-up of the rail design was built in order to practice loading the balls, the rails were found to be too fragile to withstand being rapidly loaded. If it were to be a machine loading the balls with the exact same method and without error, the rails would have been sufficient. Human error during the load process was not a risk the team was willing to take. In order to make the rails less fragile, weight would need to be added to reinforce the rails in critical spots. This would add to the empty weight of the aircraft decreasing the score. As well as the weight savings from the rails, additional weight savings occurred in the length of the landing gear. The landing gear was originally designed to accommodate six balls. Since the maximum number of balls that will be carried with the pylons is two, the landing gear was able to shrink, saving weight. A side by side comparison of the score comparison is shown in Table 11. Rail System Score Pylon System Score Score Score Empty Weight Empty Weight 2 Fastest Loading Time 10 Fastest Loading Time 10 Loading Time 15 Loading Time 10 Max Mission 3 Balls 24 Max Mission 3 Balls 24 Mission 3 Balls 6 Mission 3 Balls 2 Table 11 Drop System Score Comparison For this reason, the drop system was modified to use pylons on the bottom of the fuselage rather than a rail system. The pylons would have teeth that would be sloped on the underside, and flat on the top. The teeth would be spring loaded so that when the balls are slid onto the pylon, the teeth spring back out to hold the ball in place. A simple drawing of the pylon assembly is shown below in Figure

29 Figure 16 Drop System Pylon Dual Purpose Rudder Servo In an attempt to reduce the amount of servos used, not only were ailerons eliminated, the rudder servo was also designated to serve as the drop system servo as well. Using the dual rate function on the remote control, extra throw would be added onto the maximum rudder throw needed to operate the aircraft in normal flight. With the dual rate function turned off, maximum throw of the rudder would not deploy a ball. This would ensure that a ball was not accidentally released during normal flight. Once the dual rates switch is turned on, an extra five degrees of throw would be added onto the maximum. This extra five degrees would operate the drop system. The servo would pull on a string, which would pull an arm. The arm would pivot around a rod. The side of the arm connected to the rudder servo would be about a quarter inch long. On the opposite side, the arm would be five times as long. The arm is used to get additional throw from the rudder servo since five degrees is not much deflection. The arm allows more deflection to operate the drop system. The longer side of the arm is connected to the teeth of the pylon, once the system is activated, the string will pull the teeth inside the pylon, allowing the ball to fall off the pylon. A layout of the Drop system is shown below in Figure 17. Figure 17 Rudder Drop Layout 29

30 5.0 Detailed Design Detailed design is used to outline all of the structural and system layouts. Final performance predictions and numbers are generated using the updated weights from detailed design. Detailed design consists primarily of acquiring components and producing detailed drawings for construction. 5.1 Aircraft Dimensional Parameters Aircraft dimensional parameters define the external geometric shape of the aircraft. Table 12 shows the overall dimensions and other geometry specific characteristics. Dimensional Parameters Wingspan 72 in Wing Area in 2 Wing Chord in Length 40 in Height in Tipback Angle 10 H-Stab Span 21.6 H-Stab Chord 7.73 in H-Stab Area in V-Stab Span 9.0 in V-Stab Chord 6.87 in V-Stab Area in Table 12 Dimensional Parameters 5.2 Structural Characteristics Critical Design Elements Through our scoring analysis shown in section 3.2 we have concluded that in order to design a competitive plane the structure will have to be as light as possible while maintaining adequate structural stiffness with enough volume to accommodate the large payload size for mission 2. The team hopes to accomplish this by focusing on these key parameters during the design process: Weight: Focus on preventing the overbuilding of members by designing the aircraft to meet the structural requirements outlined in the proceeding sections. This is accomplished by introducing the practice of performing light calculations before more rigorous math or FEM software is used. This allows us to re-design the components until they fail where we want them to. Multi-Purpose Structures and Continuous Parts: Maintain a strict policy of combining structural members whenever possible. This allows us to simplify the design and lower part count which permits the use of less intensive analysis, reduces the amount of adhesive, and creates the 30

31 opportunity to design continuous and unbroken parts where failure at a joint is not a concern. Simplifying the design will also have positive effects on manufacturability as well as reparability. Manufacturability: Manufacturability is enhanced through the reduction of part count, the proper selection of building materials, and designs which facilitate quick build times. Our goal is to design uncomplicated parts that can be manufactured using either a laser cutter or a foam cutter. Designing an uncomplicated structure facilitates fast build times; while parts that can be reprinted quickly and easily makes it possible to modify the design if a critical anomaly is identified during testing or to quickly repair the plane in the event of a crash. Material Selection: Structural stiffness is critical to obtaining the predicted aerodynamic performance of the aircraft, thus it is critical to achieving a high mission score. Materials are selected based on their strength to weight ratio and their density when moment of inertia or volume is a concern Load Paths The layout of the aircraft must be able to accommodate the internally stored payload for mission 2 as well as the external payload mounts for mission 3. The layout must also give quick and easy access to the propulsion and control systems. These constraints require us to have an internal payload bay greater than 5.5 x 4.5 x 10 and an external mounting provision capable of storing and dropping at least one 12 inch circumference wiffle ball. The primary structural members are a balsa and plywood semi-monocoque wing with main and rear spars that are bonded to the Figure 18 - Structural Load Paths payload floor. These are attached to a semi-monocoque floor which holds the mission 2 payload. Main 31

32 landing gear and the wing spar also tie into the floor structure. A balsa frame fuselage and a balsa truss tail boom are detachable for transport. The load paths are shown in Figure 18 with a top and isometric view to better identify the fuselage, wing, and tail load paths Material Selection The materials were chosen with the ultimate goal of increasing structural stiffness and reducing weight. Balsa makes up a considerable amount of the structure of the plane because of its high specific strength and its ability to be used in a laser cutter at thicknesses up to ¼ inch. Where stiffness is critical and space is limited we have substituted balsa for plywood. Plywood is used in the construction of the inboard portion of the spar, the main structural component of the payload floor, the firewall, the fuselage-tail bulkheads, and the landing gear trunnions. Polystyrene foam was used in the empennage because of its light weight and its ability to be manufactured and repaired extremely quickly. Foam was also used to buttress the stringers that run across the fuselage and tail boom frames. 3D printed nylon was used to construct the materials for the payload drop mechanism Wing Design The wing is designed as a semi-monocoque stressed skin beam with a main spar at 25% chord, a rear spar at 80% chord, and a 1 / 16 inch skin placed over the leading and trailing edges with a layer of Microlight covering the entire structure. The main spar and leading edge skin are the main structural members of the wing and are designed to support the longitudinal bending moments experienced during flight. The rear spar resists torsion and provides stability to the aft portion of the wing. 1 / 8 inch thick balsa ribs are placed at varying distances along the span to prevent buckling of the skin. The rib spacing is calculated from the Euler column bucking formula 12. A 3 / 16 inch thick rib is used at the interface between the wing and the fuselage. This rib will be bonded to the frame structure of the fuselage and serves to transfer the shear and bending loads from the skin to the fuselage frame. Our initial design was a fully covered wing with a 1 / 32 inch thick balsa skin, however, the twist was too great due to the moments caused by the distance between the center of lift and the center of gravity of the wing. We removed the skin in the middle of the chord and placed it at the leading and trailing edge sections, giving a Figure 19 - Wing Design total skin thickness of 1 / 16 inch. The spar is designed in sections to allow for manufacture with a laser cutter. The sections are then bonded together in a square cutout pattern to allow for a larger adhesive surface area. To increase 32

33 structural rigidity the inboard portion of the spar is constructed from 1 / 16 inch plywood and is bonded to the main structural supports of the fuselage and payload floor. The spars shift to balsa 11.5 outboard of the centerline of the aircraft Fuselage and Cargo Bay Design The fuselage is designed to create the forward geometry of the airplane and contains the most critical structural elements after the main wing. The fuselage consists of a 1 / 16 inch thick balsa frame covered in Microlight with 1 / 16 inch stringers that are placed along the profile to stabilize the geometry and to transfer torsional loads across the structure. The fuselage is modeled from a NACA 0015 airfoil which serves several distinct advantages over more traditional rectangular or cylindrical designs. The naturally streamlined profile provides extremely low drag while the large volume allows us to scale the design to fit the mission 2 payload while also providing ample space for the control and propulsion systems. Additionally, the airfoil profile provides a continuous arching shape that is inherently strong. The fuselage design and components are depicted in Figure 20. The wing spars run through the cargo bay and are bonded to two plywood supports creating a fixed wing structure. The cargo bay is then bonded to the two sides of the fuselage frame. We chose to design the cargo bay as a major structural component of the aircraft because it allows us to create one structural member that is both extremely strong and extremely light weight, while at the same time this design reduces complexity and part count. We have added a latched cargo door to the top of the fuselage in order to access the internal structure quickly and easily. It is designed as a balsa frame with a tape hinge and is opened via a 3-D printed latch that locks into the closed position by a spring. Figure 20 Fuselage Components 33

34 5.2.6 Tail and Landing Gear Design Because the competition is in Tucson this year the airplane design must satisfy the constraints that shipping has on size. We decided on designing a removable tail as opposed to the more common removable wing because of the strength that is added to the wing by bonding the spar to the payload floor section and maintaining a continuous skin. Because the loads that are created by the tail during flight can be broken into shear forces and bending Figure 21 - Tail Boom and Empennage moments we chose a structure that was a combination of a truss and a hollow rod. The tail assembly is shown in Figure 21. The truss structure consists of the sides of the tail boom and is designed to transfer the forces created by the horizontal and vertical tails to the fuselage frame and cargo floor. The cross beams of the tail boom stabilize the truss structure and resist the torsional loads created by the rudder. The tail section attaches to the fuselage via a plywood bulkhead outfitted with dowels to transfer the torsional loads to the fuselage. The tail is fixed to the fuselage by a series of clamps and pins for flight testing but will be bonded together after transport to the competition in order to decrease the weight of the airplane. The tail is designed to create a 7.8 angle between the horizontal and the sweep of the boom to minimize separation drag. The angle was chosen as a value under 15 that would meet the horizontal and vertical stabilizers at the required value for L t. The sides of the tail boom also sweep in at an angle of 8.4 for the same reason. The horizontal and vertical tails consist of polystyrene foam with plywood spars that run through the centers of lift of their respective geometries. These spars Figure 22 Tail Boom Interface are bonded to the tail boom via a foam balsa interface shown in Figure 22. Along the aft portion of the horizontal and vertical stabilizers is another plywood spar that serves to strengthen the tape hinge that attaches to the rudder and elevator. The horizontal and vertical stabilizers are then boned to a balsa and plywood bulkhead that serves as the interface between the tail and the tail boom. The rear landing gear trunnion also bonds to this structure, eliminating the need for additional structural support elsewhere. The main landing gear consists of a plywood trunnion with a 30 34

35 gauge steel wire that meets the foam 2 tires 8.3 outboard of the centerline of the airplane. The landing gear is designed per requirements in Raymer 4, and allows the airplane to sit on the ground at an angle of 10 ; just under the stall angle of the main wing. The main gear is swept forward at an angle of 30 which allows the steel wire to deform in the event of a hard landing while transferring the bending loads through the trunnion and into the cargo floor. Figure 23 - Landing Gear Attachment Critical Load Case The critical load case for the aircraft was chosen as a 3g turn during mission 2. This is because the loads experienced during this mission are more than twice the loads experienced during missions 1 and 3. The maximum lift is calculated assuming a lb. gross weight. A 1g gust factor is then included. The net lift is then multiplied by a safety factor of 1.5 to give a maximum design lift of lb, equivalent to a 6g turn during mission 2. The maximum design lift is then used to calculate lift per unit span using Schrenk s approximation for span wise lift distribution given in Raymer 4. The equation obtained from this exercise is then integrated over the half span to obtain the shear force distribution. Integrating again gives the bending moment distribution. The results are shown in figure 24 with a maximum shear force of 21.4 lb. and a maximum bending moment of 356 in-lb. Because the landing gear attaches to the fuselage and not the wing, the loads during the 6g turn are equivalent to a 6g landing on one wheel located 8.3 inch outboard of the centerline of the aircraft. Because the location of the landing gear is dictated by ground stability and the factor of safety is above the recommended levels, we will use the 6g turn to evaluate all structural members other than the landing gear itself. Structural failure before this 6g load is not permissible. 35

36 5.2.8 Stress Analysis and Optimization The individual components of the aircraft were analyzed at the critical flight condition outlined in section The stress analysis and optimization process involves first determining the applied loads and reactions and the nature of the stress that the part will experience. Next involves a quick analysis using basic strength of materials concepts Hibbler 12. At this point it is determined if the part is overbuilt or underbuilt. Necessary changes are made and then analyzed using more rigorous methods outlined in Allen 13. From this point it is permissible to utilize FEA methods and software in order to validate predictions. All analysis is performed assuming isotropic materials and small deflection theory. The inherent disadvantage to this is that wood is an orthotropic material with lateral and shear strengths that are much lower than its axial properties. To account for this error shear stresses are calculated in the critical components and compared it to the predicted material strengths However, static testing of the airframe will be completed to validate our analysis. The wing and spar stresses are shown relative to their position on the chord in Figure

37 Moment (in-lb) Shear Force (lb) Lift Distribution (lb/in) Distance along span (in) Distance along span (in) Distance along span (in) Figure 24 Lift, Shear, and Moment Distribution 37

38 5.3 Systems and Sub-Systems Selection Subsystems selected are covered through the report and in section 5.4. Standard commercially available RC components were selected for this aircraft. Futaba 31114M servos were selected for their weight to torque. Radio components selected were the Spektrum TM DX8 transmitter and accompanying AR6210 receiver. Motor and ESC components were specked from Electrifly TM. The Rimfire TM.15 and Electrifly TM 35A ESC have proven to be a reliable system that is very durable. Battery packs were sourced presoldered using ELITE 1500 cells. All other hardware components were either hand built or sourced from the local hobby supplier. 5.4 Aircraft Weight and Balance Finalized aircraft specifications were tabulated for performance evaluations. An aircraft weight was tracked from the beginning of preliminary design. A build-up method was used to track the weight of each component. The structures design group was held to a strict weight allotment. Structural weight allowance was found by subtracting know component weights from the target weight set during scoring analysis. Aircraft weight build-up is show in Table 13. Using the weight build-up, the S&C group was able to track CG locations and properly size the tail surfaces. Each flight mission has a specific CG location for optimal performance. Weight and Balance figures are shown in % MAC and % Static Margin for clarity. See Table 14. Weight Buildup Component Weight (oz) Motor (Rimfire.15) 4.3 Battery Pack Rx Battery.5 Rx (Spektrum AR6210).46 ESC (Electrifly 35) 1.5 Servos (2 Futaba 3114s) 1.0 Propeller (APC 10x6, 8x8).75 Landing Gear 1.15 Fuselage Structure 3.84 Wing Structure 5.92 Tail Sufaces 1.12 Microlight Covering.5 Adhesive 1.6 TOTAL AIRCRAFT 34 (2.125 lbs) Table 13 Component Weight Buildup Aircraft Weight & Balance Configuration Weight CG % MAC Static Margin Mission lb 36 % 9.5 % Mission lb 36 % 9.3 % Mission lb 35 % 10.3 % Table 14 Aircraft Weight and Balance 38

39 5.5 Flight Performance Parameters Final predictions for aircraft performance were compiled during the detailed design phase. These values will be compared with testing to validate their performance. Propulsion system parameters can be seen in Table 15. Flight performance parameters define all of the critical flying speeds, take-off distances, and load factors. Climb rates and turn radii are also included in Table 16. Propulsion System Parameters M1 M2 M3 Motor Rimfire.15 Battery 14 Cell Elite 1500 Voltage 15.4 (14.0 after losses) Propeller APC 8x8 APC 10x6 APC 8x8 Max Current amps 16.5 Endurance 5.5 min 3.2 min 5.5 min T / W Static Thrust 1.75 lb 4.0 lb 1.75 lb Table 15 Propulsion Parameters Flight Performance Parameters Empty Weight lbs RAC Vs 36 ft/sec Vlof 43.2 ft/sec Vc 90 ft/sec TOFL (Mission 1/3) 36 ft TOFL (Mission 2) 58 ft ALD 120 ft n max Turn Radius (M1/M3) Turn Radius (M2) ROC (M1/M3) ROC (M2) 3.0 G 43 ft 89 ft 28 ft/sec 22 ft/sec Table 16 Flight Performance Parameters 39

40 5.6 Mission Performance Mission performance predictions were compiled using the final flight performance predictions. The total score is shown in Table 17. Mission Performance Parameters Empty Weight lbs Number of Servos* 3 Ground Loading Time 10 sec Mission 1 Laps 9 Mission 2 Time 72 sec Mission 3 Balls 1 Total Mission Score 188 *Number of Servos includes ESC Table 17 Mission Performance 40

41

42 42

43 43

44 44

45 6.0 Manufacturing Plan Manufacturability was kept under consideration thought the entire design process. Care was taken to avoid unnecessary complexity. Because the majority of the team members had previous experience with building both testing models and small DBF type aircraft, there was a very good understanding of the materials and facilities available for use through university labs and local suppliers. However, even with the group s collective knowledge, it was nonetheless important to validate the selected building materials. Composites, wood, and foam were all considered building materials. Materials were validated by the structures group and then selected for various applications on the aircraft. 6.1 Selected Manufacturing Methods A balsa build up construction method was selected for the main airframe and would be covered with a lightweight MicroLight TM style covering. Thin spruce ply-wood was selected for high stress areas where point loads would be carried. Foam was deemed acceptable for lower stress areas where a high surface area needed to be covered or high moments of inertia were required. Composites were considered for use in the spar and other high stress areas. However, it was determined through analysis that the composite layup would outweigh a similar balsa setup for a given strength requirement. In many cases, composites were just too heavy for use in such a small lightweight application as a competitive DBF aircraft. 6.2 Aircraft Manufacturing Process Tooling is the most important part of the building process. In order to achieve a high quality resulting product as well as reduce material, tooling was designed for almost every aspect of the aircraft. Wing tooling allowed the dihedral and incidence to be set accurately. Tooling for the main structure was laser cut from basswood and designed to be re-useable for multiple aircraft and the wind tunnel test article; figure 25. Figure 25 Wing Tooling

46 Molds were cut out of polystyrene Foamular150 in order to help form the wing leading edge skins. Skins are soaked in an Ammonia based glass cleaner and then pressed in the molds until dry. This is a proven process that has been used for multiple designs at Wichita State University. Figure 26 shows the wing skins being formed for the wind tunnel test article. Figure 26 Leading Edge Skin Forming 6.3 Manufacturing Milestones Manufacturing resources were allocated for drop mechanism test beds, structural test articles, the wind tunnel test article, prototype, and production aircraft. See Figure 27. Figure 27 Manufacturing Schedule 46