Milestone 1: Aircraft Design

Transcription

1 Iowa State University AerE 294X/AerE 494X Make to Innovate Milestone 1: Aircraft Design Author(s): Joshua Buettner Joseph Cairo Michael Londergan Jose Montesinos Theodore Permula Samuel Ruhlin Project: Cardinal Flight Team: Aerodynamics and Structure Team Role: Team Leader Team Member Team Member Team Member Team Member Team Member Team Member Faculty Adviser: Cdr. Daniel Buhr November 10, 2017

2 Contents Abstract 2 1 Introduction 4 2 Background Tasks Deliverable Problem Identification 6 4 Problem Solution Tilting Mechanism Fuselage Wing Tail Design Theory Structures Aerodynamics Design Verification Structures Aerodynamics Discussion 15 8 Conclusion 15 A Assembly Photos 16 B Linkage Calculator 17 C Aerodynamic and Wing Force Calculator 18 1

3 Abstract Cardinal Flight has set out to design, build, and fly a tilt-wing aircraft leveraging vertical launching capabilities and efficient forward flight. The Aerodynamics and Structures team is tasked with the aerodynamic design of the aircraft and the design of all the structural components of the aircraft. The design started with relatively few constraints and analysis was done to incorporate all of the desired features for the aircraft. Design changes were made throughout the process as more components were modeled and were analyzed in SolidWorks. The deliverable for this milestone includes the report that consists of the design considerations/decisions, a description of the design, and a complete SolidWorks model. This report also details problems that arose during the process and makes recommendations for improvements to future design work. Figure 1: Aircraft in forward flight configuration 2

4 Figure 2: Aircraft in VTOL flight configuration 3

5 1 Introduction Milestone 1 of Cardinal Flight, Aerodynamics and Structures team, is to create a complete design of the project aircraft that will be constructed in the Spring semester. This includes designing and modeling the entire aircraft structure, including the tilting mechanism for the wing/tail. These components were then inserted into a larger assembly to model the entire aircraft. In addition, structural and aerodynamic analysis was performed using classical methods to ensure that flight critical parts were properly sized while reducing the weight as much as possible. By modeling the assembly, the team was also able to provided weight estimates to the Electrical team for motor and electrical system sizing. All CAD modeling was done in SolidWorks. 2 Background For the academic year Cardinal Flight has set out to design and build an aircraft similar to NASA s Greased Lightning aircraft. The aircraft will use a tilting wing and stabilator to transition between vertical and forward flight. The project seeks to develop a design that can capitalize on the vertical takeoff or landing (VTOL) capability while adding in the efficiency of fixed-wing forward flight. This will allow the aircraft to carry a wide range of payloads and support them through all flight regimes. 2.1 Tasks This milestone was comprised of the following tasks: Sept Preliminary Tilting Mechanism Design Sept Final Tilting Mechanism Design Oct Model Wing in SolidWorks Model wing tilt mechanism Model internal components Oct Model fuselage in SolidWorks Model internal components Model tail 4

6 Oct Complete SolidWorks Assembly Nov Complete Milestone Report 2.2 Deliverable Milestone Report Design Considerations and Decisions Design Description Full SolidWorks Model 5

7 3 Problem Identification The Aerodynamics and Structures team s goal for this milestone was to take a conceptual idea that was developed the previous academic year and turn it into a functional and safe design. There are a few technical issues that had to be overcome to accomplish the aircraft design. First the team needed to determine if the project was feasible, followed by designing the aircraft, and then designing all the internal components for the aircraft. The aircraft design process began with a constraint of a wingspan of six feet. This was due to transport concerns if the wing was larger six feet. A root chord of eighteen inches, and a tip chord of ten inches for the wing was decided on after preliminary lift calculations. The desired wing design would include a wing sweep to be determined in the design phase to manipulate the center of lift and the center of gravity. The team needed to evaluate the aerodynamic properties of the chosen Clark Y airfoil to determine if the wing will produce sufficient lift for the designed structure. The structure of the aircraft also needed analysis to ensure a capable design. Force calculations will be completed for major structural components such as the wing spar and tail boom. In addition, an analysis will be performed on the wing tilt mechanism to ensure the servo is capable of the handling the load from the wing. 4 Problem Solution 4.1 Tilting Mechanism To determine whether or not the project was feasible, the team first examined the tilting mechanism for the wing and tail to determine how it could be accomplished. Several solutions were explored. The first and most desirable solution was to use a linear actuator for the wing. This was the most desirable solution due to the load bearing capability of the linear actuator and once power was removed from the actuator, it would remain at it s position and would be able to handle very high loads. Weight and space constraints lead the team to move away from this plan in addition to the need for it s own power supply. The linear actuator also moved slower than desired for the transition from vertical to forward flight The next idea was to use a servo gearbox. This would allow us to place a normal servo into a gearing mechanism to greatly increase the torque produced by the servo. This solution was better because it allowed the wing to transition quicker than the 6

8 Milestone 1: Aircraft Design linear actuator and it was lighter than the linear actuator. As the team continued the aircraft design, it became apparent that the aircraft structure weight needed to be reduced. This meant that the servo gearbox was too heavy and another solution was needed. Figure 3: Initial wing tilt design with servo The next solution was a high quality, metal geared servo. This is the most light weight solution that the team was able to find. It gives the aircraft the ability to transition quickly and with a more compact footprint for the mechanism inside the aircraft. The downside to this is that the required torque had to more carefully calculated to determine if the servo was sufficient. To determine the feasibility of different linkage combinations for the wing and tail, team members developed a simple MATLAB program that would take geometrical inputs and output the available degrees of rotation for the system. A feasible linkage was designed to be placed into the SolidWorks assembly. After installing the system into the full assembly, it was determined that the linkage would not work in its current position. The servo had been mounted at the forward edge of the payload plate which is 1 aft of the leading edge (LE) of the wing. This led to an evaluation of other possible linkage combinations. If no other suitable solution could be found, the fuselage would need to be modified to accommodate the tilting mechanism. This preliminary design can be seen in Figure 3. Fortunately, a suitable solution was found that did not necessitate the need to change the design of the fuselage. The servo was relocated to 7 aft of the wing LE. It was also placed onto the tail boom that extends into the fuselage. This will provide a solid mounting platform and moves the servo 1 up. These two changes combined 7

9 Milestone 1: Aircraft Design Figure 4: Final wing tilt design with servo moved allow the mechanism to use a shorter linkage that is capable of fitting in the fuselage and providing a full 90 of rotation. This satisfies the requirement for the wing tilting mechanism. The final design can be seen in Figure 4. If it is determined that the selected servo is not capable of sustaining the load from the wing, an additional servo will be added to the system. The tail tilting proved to be a simpler problem due to it being mounted on the tail boom with no space constraints. The MATLAB program Appendix B was utilized to determine the best linkage size for the system. The servo is located approximately 14 forward of the stabilator spar, which is the pivot point. The stabilator design was chosen to simplify the design and reduce the need for additional servos to manipulate elevators. The linkage between the stabilator and servo arm is 15. This system provides rotation to 90 vertical and -45. This is more than adequate for use as an elevator and tilting mechanism. The design is shown in Figure Fuselage After several iterations for the fuselage design, the team decided to use standard aircraft construction methods. The fuselage is a semi-monocoque design using bulkheads to define the shape and stringers/longerons to increase the strength of the fuselage. The bulkheads made from carbon fiber plate. The stringers/longerons are carbon fiber pultruded rods that will fit into cutouts made in the bulkheads. The fuselage will be covered in Monokote with a removable fiberglass nose cone. An additional structural component of the fuselage is a carbon fiber and foam sand- 8

10 Figure 5: Tail tilting design wich plate. This is located near the center of the aircraft. It provides an area to attach the payload pod, but it also is used to mount the brackets to hold the wing and the attach point for the tail boom. The fuselage is designed so that the electronic components can be located in the forward part of the fuselage to assist in center of gravity manipulation. This area will be accessed through a removable fiberglass nose cone. In addition since the brackets to hold the wing are attached to the payload plate, the brackets can be easily removed which translates to easily removed wings. This capability is essential for transporting an aircraft of this size. 4.3 Wing The aircraft s wing is a more complicated structure than previous Cardinal Flight Designs. The main reason being due to the wing being tilted in flight. This required some analysis for spar strength and also determining the best place to pivot the wing at. To do this, the neutral point (NP) of the aircraft was calculated using the MATLAB program which can be found in Appendix C. The NP of the aircraft is the point on the aircraft where the change in moment due to a change in angle of attack (AOA) is zero. If the center of gravity (CG) of the aircraft is aft of this point, the aircraft will be uncontrollable without a computer flight control system. Since the CG of the plane needs to be aft of the wing when 9

11 Figure 6: Internal structure of wing it is in VTOL mode, the pivot point on the wing needed to be positioned so that it was forward of the NP, but also forward of the CG when in VTOL operations. To do this, the team decided to pivot the wing nine inches aft of the leading edge of the wing. This would meet all of the conditions for CG and NP locations. This aircraft will be using multiple motors on the wing for both VTOL and forward flight modes. This required evaluation of the structure and a design that could support the eight motors distributed across the wing. This included using two spars in the wing. The main spar is located at the quarter chord location throughout the entire wing. The main spar in each wing is joined in the center section to make a one piece spar. This spar is also what the motor mounts will be attached to in addition to a pultruded rod that spans the leading edge of the wing. There was also a need for structure at the pivot point for the wing. To do this, a second spar was added that extends out to the middle of the wing. This adds additional strength to the wing in both VTOL and forward flight modes. The second spar is located 9 aft of the wing LE. 4.4 Tail Another step taken in the design process involved the sizing and modeling of the stabilator, the vertical stabilizer, and the rudder. The stabilator design was chosen in order to simplify the horizontal tail. Since the entire control surface would be rotating for the hovering configuration, it would be much more complex to add a functioning elevator within the horizontal stabilizer and it would increase weight. The stabilator was designed using carbon fiber ribs with constant chord. There is a 0.5 carbon fiber tube running down its quarter chord. In order to have the control surface rotate, two bearing housings will be mounted to the sides of the rear fuselage. 10

12 Figure 7: Stabilator These will hold the stabilator spar in place while a servo can rotate one of the center ribs via a control horn thereby rotating the stabilator. The vertical stabilizer is designed to be constructed of tapered carbon fiber ribs using a NACA0012 airfoil with a carbon fiber tube spar located in the center of the rib. The rudder will also be constructed using carbon fiber ribs. Figure 8: Vertical Stabilizer 11

13 5 Design Theory 5.1 Structures To evaluate if the spars were of sufficient strength, bending stress calculations were used. An ultimate stress of 500 ksi was used and the team worked backwards to determine the load that would be required to reach that point. These equations are over simplified for composites, but a reasonable estimate of the load and factor of safety can be found. σ b = My I (1) M = Σ(F d) (2) I = π(d4 outer d 4 inner) 64 (3) 5.2 Aerodynamics The airfoil used for the wing is a Clark Y and the tail uses a symmetrical NACA The Clark Y airfoil is a common airfoil used for many purposes. The airfoil performs well at the speed and altitude the project aircraft will be operating at. Analysis was done on the wing to ensure that sufficient lift will be created during forward flight. Lift will be calculated using Equation 4. L = 1 2 ρv 2 C L S (4) To determine the tail specifications, the following method was used. V h = S hl h Sc (5) In Equation 5, V h represents the horizontal tail volume coefficient, S h = horizontal tail area, l h = the horizontal tail moment arm, S = wing area, and c = average wing chord. V h for a stable aircraft should fall between the values of 0.30 and The 12

14 value of 0.60 was chosen while designing the stabilator to ensure ample pitch control. Next, the vertical tail volume coefficient, V v, was determined using Equation 6. V v = S vl v Sb (6) V v, in Equation 6, represents the vertical tail volume coefficient, S v = vertical tail area, l v = the vertical tail moment arm, S = wing area, and b = wing span. V v should fall between the values of 0.02 and Due to problems in the past with insufficient size of control surfaces, the team chose the value of 0.05 for the vertical tail volume coefficient. 6 Design Verification 6.1 Structures The main spar is a standard modulus carbon fiber tube with an outer diameter (OD) of 1 and an inner diameter (ID) of The second spar is a high modulus carbon fiber tube with an OD of 0.5 and an inner diameter of For the main spar it was determined that the spar would fail when subjected to a moment of 33.6E3 lb-in. If the moment is taken about the wing tip, the required force to get that moment would be 930 lbs. With the planned aircraft weight of 15 lbs and adding 10 lbs for the thrust from the motors, the factor of safety for the main spar is 37. Since the calculations have been simplified with respect to composite analysis, this factor of safety is acceptable and excessive for the needs of the structure. The second spar was evaluated similarly and provided a factor of safety of 11. Again, these methods are simplified and the strength of the spars is likely to be much greater due to the layup of the carbon fiber plies in the spar. 6.2 Aerodynamics The team determined that the airfoil will be capable of creating at 16 lbs of lift at an airspeed of 30 mph and an AOA of 5 degrees. As the speed is increased in forward flight, the AOA could be reduced while maintaining sufficient lift. This calculated was the determining factor in sizing the entire aircraft and is how the other Cardinal 13

15 Flight teams worked on sizing the electrical/propulsions systems and the available payloads. Using the methods for sizing the tail, it was determined that the vertical stabilizer would be 12 tall. The chord of the vertical stabilizer ribs starts at 12 and tapers to 9.5 at the top. The stabilator uses a constant chord of 10 with a 30 span. 14

16 7 Discussion The milestone was a success. The team was able to take the project concept and create a complete design that will be constructed in the Spring. There were a few challenges that necessitated changes to certain components or assemblies throughout the design process. All the changes led to a better design that will also be easier to manufacture. As is the case with most aerospace designs, the design continually struggled with weight. The design materials have been used in previous Cardinal Flight projects, so the team was able to take that knowledge and focus more on the design of the structure instead of selecting materials. As it stands, the aircraft structure weighs six pounds. This estimate is from Solid- Works with information from the commercial components that will be used in the aircraft. A future goal will be to explore other ways to reduce weight in the design to maximize the payload capacity. This effort to reduce weight also created challenges, as stated previously, in the tilting mechanisms and the components that could be used to solve that problem. A recommendation to future design teams is to incorporate a SolidWorks course in the beginning of the semester to get all the team members up to speed and ensure a smooth design phase. The team was productive, but tasks could have been accomplished faster if team members had been given some baseline knowledge. 8 Conclusion Overall, the design phase went smoothly. The most problematic area was the design of the tilting mechanism, but this challenge was expected prior to beginning the project. After this, the only big challenges were working through the many parts that needed to be modeled, which took time. By using previous project knowledge, material selection was greatly simplified and the team was able to focus mainly on the design of the aircraft. SolidWorks drawings and a material list can now be created in preparation for the build phase, which will begin in the Spring 2018 semester. 15

17 A Assembly Photos Figure 9: Assembly in forward flight configuration without skins 16

18 Figure 10: Aileron structure Figure 11: Motor Mount B Linkage Calculator MATLAB program to calculate different linkage situations: LinkageCode.m 1 c l e a r, c l c 2 3 R1 = input ( Enter servo arm length : ) ; 4 r1 = input ( Enter servo r a d i u s : ) ; 5 R2 = input ( Enter t a i l /wing length : ) ; 6 r2 = input ( Enter t a i l /wing r a d i u s : ) ; 7 H = input ( Enter height d i f f e r e n c e : ) ; 8 D = input ( Enter d i s t a n c e between r o t a t i o n p o i n t s : ) ; 9 P = input ( Enter l i n k a g e length : ) ; f o r theta1 = 0 : ; 17

19 12 theta = theta1 pi /180; 13 phi = theta ; p s i = acos ( (D+R2 cos ( phi ) R1 cos ( theta ) ) /P) ; 16 f p r i n t f ( \n At %f d e grees p s i i s %f, theta1, p s i ) ; 17 dx = D R1 cos ( theta )+R2 cos ( phi ) ; f o r x = R1 cos ( theta ) :. 1 : D+(R2 cos ( phi ) ) ; 20 y = R1 s i n ( theta )+(0 R1 cos ( theta ) ) ( s i n ( p s i ) ) /(P dx ) + x s i n ( p s i ) /(P dx ) ; i f s q r t ( ( x ˆ2)+(y ˆ2) ) <= r1 ; 23 f p r i n t f ( \n %f, theta1 ) ; 24 break e l s e i f s q r t ( ( (D x ) ˆ2) +((y H) ˆ2) ) <= r2 ; 27 f p r i n t f ( \n %f g, phi ) ; 28 break 29 end 30 end 31 end C Aerodynamic and Wing Force Calculator MATLAB program to calculate neutral point and force on the wing for a given pivot point: wing.m 1 clc, c l e a r a l l 2 3 [ dens, a, v i s c ] = StandardAtmosphere ( 1000 ) ; 4 5 V = ; % kts to m/ s 6 7 r o o t c h o r d = 1 8 ; %in 8 f u s e w i d t h = 6 ; %in 9 t i p c h o r d = 1 0 ; %in 10 wing span = 7 2 ; %in 11 s w e e p d i s t = 1 5 ; %in This i s the d i s t a n c e from the root LE to the t i p LE 18

20 12 p i v o t d i s t = 9 ; % in % Find LE and TE l i n e formula f o r swept wing 15 l e b = 0 ; 16 l e a = ( s w e e p d i s t l e b ) / ( ( wing span f u s e w i d t h ) 0. 5 ) ; 17 t e b = r o o t c h o r d ; 18 t e a = ( ( s w e e p d i s t + t i p c h o r d ) t e b ) / ( ( wing span f u s e w i d t h ) 0. 5 ) ; % Find h e l p e r l i n e formula 21 mac b0 = t i p c h o r d ; 22 mac a0 = ( ( s w e e p d i s t + t i p c h o r d + r o o t c h o r d ) mac b0 ) / ( ( wing span f u s e w i d t h ) 0. 5 ) ; 23 mac b1 = r o o t c h o r d + t i p c h o r d ; 24 mac a1 = ( ( s w e e p d i s t r o o t c h o r d ) mac b1 ) / ( ( wing span f u s e w i d t h ) 0. 5 ) ; % Determine MAC using i n t e r s e c t i o n o f h e l p e r l i n e s 27 mac x = ( mac b1 mac b0 ) / ( mac a0 mac a1 ) ; % Compute MAC i n t e r s e c t i o n with LE and TE 30 le mac y = l e a mac x + l e b ; 31 te mac y = t e a mac x + t e b ; %C a l c u l a t e mac 34 mac = te mac y le mac y ; % Get the wing area and area a f t o f the pivot point 37 [ t o t a r e a, a r e a a f t ] = wingarea ( fuse width, root chord, tip chord, sweep dist, p i v o t d i s t, wing span ) ; % Wind Load C a l c u l a t i o n s with h o r i z o n t a l wind Multiplying wing area to move to mˆ2 40 F w = 0. 5 dens Vˆ2 ( t o t a r e a ) ; % N 41 p dyn = (F w / ( t o t a r e a ) ) ; % N/m % C a l c u l a t e fwd area and convert to mˆ2 44 fwd area m2 = ( t o t a r e a a r e a a f t ) ; % mˆ

21 46 % Convert area from in ˆ2 to mˆ2 47 area aft m2 = a r e a a f t ; %mˆ % C a l c u l a t e the t o t a l f o r c e a f t o f pivot point during v e r t i c a l and convert 50 % f o r c e from Newtons to l b s 51 a f t f o r c e = ( area aft m2 p dyn ) / ; f w d f o r c e = ( fwd area m2 p dyn ) / ; fwd mom = ( 1) f w d f o r c e ( p i v o t d i s t /2) %in l b s aft mom = a f t f o r c e ( ( ( s w e e p d i s t+t i p c h o r d ) p i v o t d i s t ) /2) wingarea.m 1 f u n c t i o n [ area1, area2 ] = wingarea ( fusewidth, rootchord, tipchord, sweepdist, pivotdist, span ) 2 %wingareas Given information about the wing, c a l c u l a t e the a r e a s 3 % S p e c i f i c a l l y, given a pivot d i s t, c a l c u l a t e the area a f t o f that 4 % point to be used in f o r c e c a l c u l a t i o n s. 5 6 x = [ fusewidth, ( span /2)+(fuseWidth /2), ( span /2)+(fuseWidth /2), fusewidth,0, ( span /2)+(fuseWidth /2), ( span /2)+(fuseWidth /2),0, fusewidth ] ; 7 y = [0, sweepdist, ( sweepdist+tipchord ), rootchord, rootchord, ( sweepdist+tipchord ), sweepdist, 0, 0 ] ; 8 9 f i g u r e ( ) 10 p l o t ( x, y ) ; 11 xlim ([ ] ) 12 ylim ([ ] ) area1 = polyarea ( x, y ) ; 15 x1 = [ ( p i v o t D i s t /( sweepdist / ( ( span /2) (fusewidth /2) ) ) )+ fusewidth, ( span /2)+(fuseWidth /2), ( span /2)+(fuseWidth /2), fusewidth,0, ( span /2)+(fuseWidth /2), ( span /2)+(fuseWidth /2), ( p i v o t D i s t /( sweepdist / ( ( span /2) (fusewidth /2) ) ) ), ( 20

22 p i v o t D i s t /( sweepdist / ( ( span /2) (fusewidth /2) ) ) )+fusewidth ] ; 16 y1 = [ pivotdist, sweepdist, ( sweepdist+tipchord ), rootchord, rootchord, ( sweepdist+tipchord ), sweepdist, pivotdist, p i v o t D i s t ] ; f i g u r e ( ) 19 p l o t ( x, y, x1, y1 ) ; 20 xlim ([ ] ) 21 ylim ([ ] ) area2 = polyarea ( x1, y1 ) ; area2 m2 = area ; %mˆ ( area2 m ) / ; end StandardAtmosphere.m 1 f u n c t i o n [ dens, a, v i s c ] = StandardAtmosphere ( a l t ) 2 %Read in a l t i t u d e and output v i s c o s i t y, density, and speed o f sound 3 %Units : 4 %a l t i t u d e as a s t r i n g in f t 5 %d e n s i t y in kg/mˆ3 6 %speed o f sound, a, in m/ s 7 8 %Convert a l t i t u d e to number in meters 9 a l t = str2num ( a l t ) ; g = ; 12 R = ; 13 gamma = 1. 4 ; 14 i f a l t < T1 = ; %From Anderson, I n t r o d u c t i o n to Flight, 5 ed, pg l a p s = 6.5e 3; 17 T = T1 + l a p s ( a l t ) ; 18 dens1 = ; 21

23 19 dens = (T/T1) ˆ (g /( l a p s R) +1) dens1 ; 20 e l s e i f a l t >= & a l t < T1 = ; 22 T = T1 ; 23 dens1 = ; 24 dens = exp( g /(R T) ( alt 11000) ) dens1 ; 25 e l s e 26 T1 = ; 27 l a p s = 3e 3; 28 T = T1 + l a p s ( a l t ) ; 29 dens1 = ; 30 dens = (T/T1) ˆ (g /( l a p s R) +1) dens1 ; 31 end 32 a = s q r t (gamma R T) ; 33 v i s c = e 6 Tˆ 1. 5 / (T+110.1) ; %Bertin, Aerodynamics f o r Engineers, 4 ed, pg 5 22