機械小黃鴨的機械機構(gòu)設(shè)計及運動仿真-機械鳥【Creo三維含10張CAD圖帶開題報告-獨家】.zip

機械小黃鴨的機械機構(gòu)設(shè)計及運動仿真-機械鳥【Creo三維含10張CAD圖帶開題報告-獨家】.zip,Creo三維含10張CAD圖帶開題報告-獨家,機械,小黃鴨,機構(gòu),設(shè)計,運動,仿真,Creo,三維,10,CAD,開題,報告,獨家 畢業(yè)論文附件材料 機械小黃鴨的機械機構(gòu)設(shè)計及運動仿真學生姓名：黃亞龍學 號：01410141Y31所在系部：機械工程系專業(yè)班級：14gb機電1班指導(dǎo)教師：曹靜（講師）日 期：二一八年五月 目 錄1 英文文獻翻譯11.1 Design and aerodynamic analysis of a flapping-wing micro aerial vehicle1Numerical analysis41.2中文翻譯252 專業(yè)閱讀書目412.1 機械設(shè)計412.2 機械原理412.3 機電一體化系統(tǒng)設(shè)計422.4 Creo2.0機械設(shè)計教程422.5 單片機原理及接口技術(shù)432.6 機械設(shè)計課程設(shè)計432.7 現(xiàn)代工程制圖442.8 機械零部件選用與設(shè)計452.9 材料力學462.10 理論力學461 英文文獻翻譯1.1 Design and aerodynamic analysis of a flapping-wing micro aerial vehicleAuthor：Bor-JangTsai Yu-ChunFuAbstract This paper presents the design and aerodynamic performance of a planar membrane wing as shape airfoil for the micro aerial vehicle. This simulation calculates the average lift force, L as the criteria weight of the flapping wing (weight must be lower than 8.78 g), to make one ultra-light, small size flapping wing MAV. In here two phases are discussed. First, the 3D aerodynamic calculation and flow field simulation of a planar membrane wing as shape airfoil for a MAV were studied. Analyzing the flapping wing under different frequencies and angles of attack, investigates the pressure distribution, the airfoil-tip vortex and the up-wash situation of the air flow. Second is to average lift force, L 8.78 g for designing weight limit of the MAV. The specifications of flapping wing MAV are 8 g gross weight, the 15 cm wingspan, and 5 cm chord length. In this vehicle, we employed the concept of four-bar linkage to design a flapping mechanism which simulates the flapping motion of a bird. The angles of upstroke and downstroke can be varied in the design. The total flapping angle is 73. The flapping frequency of wing is 25.58 Hz. The power source comes from motor with a LiH battery. A simple flight test was carried out and the result of the flight is going well. The actual flight distance is approximately 8 m, and the primary goal is achieved. By the way, we found the rigidity of tail wing is crucial and should be enhanced to prevent the flapping-wing MAV will be unable to revise if the MAV in a crooked condition and it will cause a crash.Introduction The micro aerial vehicle, in English is abbreviated as a MAV, according to the Defense Advanced Research Projects Agency (DARPA) of USA, the size of various aspects of micro aerial vehicle (MAV) is limited to 15 cm, the flying speed is 1020 m/s, the Reynolds number must be below 106. Regarding a flapping wing for a MAV, the most important issue at present is the aerodynamic performance. The Reynolds number of a MAV is about 105, this range of Reynolds number will cause laminar separation phenomenon occurred on the surfaces of the body. Moreover, since the definition of a MAV includes size limit, and the challenge of this work is to design an ultra-light and small size of a flapping wing MAV comparing all literatures 1,7,9,10,13, therefore by using very low aspect ratio of MAV to obtain enough lifting force, L. However, small aspect ratio will increase the three-dimensional effects on flow field. The MAV is small and the speed is low, the flight stability of a MAV is affected easily by the external wind shear or other disturbances. This research applied dynamic moving grid technology and analyzed a planar membrane wing under the low Reynolds number. Each pattern of the flap movement initiates a complex and unsteady flow field. Calculation of aerodynamic performance becomes crucial. To predict lifting force, L needs to solve the whole unsteady flapping flow field of a wing. Approaches of solving this are divided into two steps; first, we do the flow field simulation and analysis, second, we design and manufacture it. Regarding the literature survey, in 2000, Neff and Hummel 9 studied the two- and three-dimensional flow fields by plunging and pitching movement for NACA 0012 airfoil, they solved the Euler equation to simulate the flap and twist movement for the rectangular wing. In 2003, Tuncer and Kaya 13 made the movement of the upstroke and downstroke flap by using the two-dimensional NACA 0014 and they analyzed the reason which is thrust force, T produced and observe the overflow situation of its turbulent flow. In 2001, the Caltech, Pornsin-sirirak made a MAV 10, they used the titanium alloy wing of the xylene thin film, complete the altogether weight is 10.5 g, also fly for 518 sec successfully. In 2005, Delaware University in USA, Agrawal imitates an insect flight and they studied the multi-dimensional flapping movement and the twisting movement to simulate the hummingbird flap and but not becomes a MAV 1. In 2006, Lin, Hwu and Young reported the trust and lift of an ornithopters membrane wings with simple flapping motion on the journal 7, they revealed the lift force, L of a flexible flapping wing will increase with the increase of the flapping frequency under the corresponding flying speed. For the same flapping frequency, the flying speed can be increased by decreasing of the angle of attack with the trade of loosing some lifting force. The flapping motion generates the trust to acquire the flying speed. The flying speed and angle of attack combine to generate the lift force, L for flying. This paper is the most important reference to us. In recent, the design of precision balance and aerodynamic characteristic for micro aerial vehicle to measure lift, drag, rolling-moment, and pitching-moment of a MAV was reported by Suhariyono et al. 12, but measurement is for the fixed wing MAV only, not for flapping wing, the measurement of flapping wing is critical. Only Singh et al. 11 studied an experimental apparatus that incorporates flapping wings and measures the small amount of thrust generated by these wing motions is described. This methodology is used to measure the thrust generated by two wings at different wing pitch settings. Also, the effect of change in pitch phase during a flapping cycle is examined experimentally. Regarding the simulation, Larijani 6 proposed a nonlinear aeroelastic model for the study of flapping wing flight in the 2001, this paper conducted the Huangs 5 numerical analysis for the flapping wing MAV later. From Refs. 9 and 7 we know the three-dimensional movement of many birds flapping is used the standard NACA shape airfoil as wings, but actual flapping wing of MAV to be restricted in the volume and the weight. Its unlikely to use the NACA series of wing section. On the contrary, the most of the flapping wing for MAV, a planar membrane wing are used primarily. In order to imitating the insect flutter and the flight pattern, therefore, this investigation does take the planar membrane wing as a study target vehicle, discussing its aerodynamic characteristic and to predict average lift force, L as the criteria weight to manufacture a future MAV. The actual MAV was made by the wingspan is 15 cm, the mean chord is 5 cm, the weight is 8 g, the wing area is 75 cm2, the flapping frequency is 25.58 Hz of a flapping wing MAV.Numerical analysisNumerical methodIn the numerical simulation solves the speed and the pressure on this pattern flow field. It is an integral control volume method. In the control volume definition, each physical quantities is significant because the separation variable is the integral of control volume for the governing equation, therefore we must first take the separation of the governing equation to control volume of the flow field computation.Governing equation:(1)1gt(g)+div(urgrad)=S:On behalf of any independent physical quantity(ui,e,k):Diffusion coefficientS:Source coefficientAfter the numerical computation of the convergence condition which in the volume change rate is smaller than after each time the iteration that we give.(2)Ck=(|BPnPn|BP0P0|)(given value)Design budgetary estimatesEstimation of the MAV weightThe MAV weight (WTotal) may include a MAV main body weight (WFuselage), a wing weight (WRudder), the load weight (battery and switch or joint) (WPayload) and the power unit (motor) (WPower).Aerodynamic parameter estimatesLift coefficient:(3)CL=2LU2SThrust coefficient:(4)CT=2TU2SReduced frequency:(5)K=fcUAdvance ratio:(6)J=U2fR=Flying speed or fluid velocityWing tip speedReynolds number:(7)Re=CUt=4fR2ARDiscussion and analysis of the numerical results. Numerical simulation Geometry contour and grid establishment In order to conform to DARPAs definition of the MAV, therefore this research takes 15 cm as the wingspan length and only constructs the single wing (half wingspan) of grid. The main consideration of chord length is for hoped the induced drag is small but the wing induced drag following the lifting force, L occurs, the lifting force, L is bigger and the induced drag is also bigger. But the wing induced drag is directly related to the aspect ratio and if the aspect ratio is bigger, relatively, the induced drag will be smaller. Therefore, this research designate the aspect ratio is 3, the chord length c is 5 cm, the thickness of planar membrane wing is 0.3 mm and the rectangular shape of wing. The grid uses the non-constructive grid, the non-constructive grid is easier than the constructive grid to process the complex geometry, has the convenience to use the three-dimensional dynamic moving grid skills 4 as well. The connecting positions of wing entity and the flow passage will have the boundary layer effect, therefore the grid became dense but the entire flow passage used the dispose for the C grid, the total grid point is 854,090, shown in Fig. 1. The computational domain; the length is 32.5c, the extended is 12.5c, the height is 25c.Setting of boundary conditions The predetermined MAV flying speed is 10m/s, therefore incoming air speed is 10m/s. The outflow is an atmospheric pressure. Because of flow field assuming sliding, therefore the hypothesis of flow passage flank is the sliding boundary, then, the position of boundary will not have the boundary layer effect. The flapping angle is 30. 3.1.3. Numerical algorithm and setting of parameters The convection terns of momentum equations use different approaching principles by spatial separation variables, two principles were employed in this study, the pressure term uses staggered type of PRESTO (Pressure Staggering Option) principle. In addition to the speed-pressure field coupling uses the SIMPLE principle. For the time accuracy, the time step is carried on iterations by the two step implicit expression law (2nd order Implicit Algorithm) 3. The important parameter settings are as follow: 1. Reduced frequency K setting: the K value is 0.1 and 0.2 and 0.3, from Eq. (5), may know the actual flight of birds conversion to the flapping frequency. K=0.1, is equal to flap of 6.369 times in each second. K=0.2, is equal to flap of 12.739 times in each second. K=0.3, is equal to flap of 19.108 times in each second. 2. Angle of attack setting: designates the angles of attack is 0, 5 and 10. 3.2. Program validation by a case of three-dimensional rigid wing Based on 2004, Ref. 5, in view of aerodynamic analysis for a three-dimensional flapping wing, simulates the behavior of the NACA 2412 rigid wing flap. Case uses the same wing section and the flow field conditions. That is the NACA 2412 rectangular wing and AR is 8, and the single wing of grid was constructed, namely half wingspan is 4c (c is the chord length 3.4 cm), the is 15, the angle of attack is 0, the U is 8.6 m/s, the flap frequency is 8 Hz, 16 Hz and 24 Hz respectively, carries on the computation of dynamic flap of unsteady flow field. The grid distribution is shown in Fig. 2, and the total grid number is 641,624. Fig. 3 is a comparison of lifting coefficient in condition of unsteady state, result of lift coefficient between this research and Ref. 5 is quite close, this proves that the setting of boundary conditions and numerical model is accuracy and correct.A three-dimensional case of planar membrane wing in differentKAOA=0,K=0.1, 0.2, and 0.3Lifting force and thrust forceWhen the angle of attack is 0 and theKvalue is 0.1, 0.2 and 0.3 respectively, investigates the increasing ofKto influence on the aerodynamic forces.Fig. 4shows the comparison of lift coefficient,CLand different drag coefficient,CDbased on differentKvalues, in the lift coefficient,CLportion, the movement of flap wing starting the downstroke and arriving the center point position from the highest peak, the lift coefficient,CLelevates to the maximum value, the movement of flap wing flapping again from the center point downstroke to the perigee position, and the lift coefficient,CLfalls to the starting value. Therefore, in downstroke for the lifting force,Lis positive. Starting to upstroke, the flap flapping from the perigee to the center point position, the lift coefficient,CLfalls to the minimum value, the movement of flap wing flapping again from the center point to the peak position, the lift coefficient,CLrises to the starting value, thus the lifting force,Lis negative value in upstroke.The increase ofKcauses the profile of top and bottom oscillation amplitude for the lift coefficient,CLto become the proportional increasing, while in downstroke, the positive lift coefficient,CLbecomes the proportion to increase. WhileK=0.1, the maximum of lift coefficient,CLis 0.1. WhileK=0.2, the maximum of lift coefficient,CLis 0.2. WhileK=0.3, the maximum of lift coefficient,CLis 0.3. While in upstroke, the negative lift coefficient,CLbecomes the proportional increasing. WhileK=0.1, the smallest lift coefficient,CLis 0.1. WhileK=0.2, the smallest lift coefficient,CLis 0.2. WhileK=0.3, the smallest lift coefficient,CLis 0.3. Increase of the positive and the negative counterbalances mutually, thus theKvalue increase does not have the contribution to the average lifting force,L(equal to zero), therefore flapping like this way is unable to generate the lifting force,L.Moreover, in the drag coefficient,CDportion, while theKincreases, the drag coefficient,CDhas the big variation only when the flap starts flapping. WhileK=0.1, the biggest drag coefficient,CDis 0.0125. WhileK=0.2, the biggest drag coefficient,CDis 0.014. WhileK=0.3, the biggest drag coefficient,CDis 0.015. The drag coefficient,CDreduces relatively when theKvalue increases, after the first flap cycle, no matter howKvalue is, both in downstroke and in upstroke will not have big change, the mean drag coefficient,CDis 0.018. As a result, while the angle of attack is 0, the increase ofKvalue does not have a quite big contribution to the average thrust coefficient.Wing tip vortexIn order to ensure the accuracy, the second period of flap cycle in numerical calculation was selected to observe, it separately picks six points of time period in the cycle to observe.Fig. 5shows the t/T=0/6t/T=5/6 are in order.Figs. 6 and 7show the velocity vector diagrams forK=0.1andK=0.3respectively, at the position of 1/4 chord length observes the wing tip vortex. While thet/T=0starting downstroke, then curls up the counterclockwise rotation of the wing tip vortex, the strong turbulent flow causes the low pressure region for the upper wing surface, therefore it may bring the upward lifting force,Lfor the plate wing. While thet/T=3/6in the perigee position of downstroke, instantaneously, the turbulent flow can be absorbed because of the big reacting force. While thet/T=4/6starting upstroke, then curls up the clockwise rotation of the wing tip vortex, the strong turbulent flow causes of the low pressure region for lower wing surface, therefore the negative lifting force,Lis not favor for the MAV flight.WhileK=0.1, no matter how the downstroke or upstroke is, the wing tip vortex appears smooth. WhileK=0.3, the wing tip vortex can be seen obviously and the average vortex velocity is 8.02 m/s for the wing tip. As a result of theKincrease can cause the maximum vortex velocity increasing quickly for the wing tip, wing tip vortex became obvious, it affects the pressure between upper and lower surfaces of airfoil, and influences on lifting force,Land thrust force,Tas well. Regardless of the increasing ofK, the upstroke and downstroke have the same clockwise and counterclockwise strength of the vortex, therefore the positive and the negative of lifting force,Lis mutually offset. This causes the average lifting force equal to zero. This result verifies thatCLandCDof differentKat AOA=0 as our expectation.A three-dimensional case of planar membrane wing in different angle of attack K=0.3,AOA=0, 5 and 10Lifting force and thrust forceK=0.3, AOA=0, 5 and 10, investigates the increasing ofKto influence on the lift coefficient,CLand the drag coefficient,CD.Fig. 8is the comparison of the lift coefficient,CLand the drag coefficient,CDunder the different angle of attack, so the increasing angle of attack conducive to favor the lifting force,Land the thrust force,Tgeneration, while in downstroke the positive lift coefficient,CLbecomes the proportion to increase. While AOA=0, the maximum lift coefficient,CLis 0.3. While AOA=5, the maximum lift coefficient,CLis 0.5. While AOA=10, the maximum lift coefficient,CLis 0.7. While in upstroke, the negative lift coefficient,CLbecomes the proportional reducing actually. While the AOA=0, the smallest lift coefficient,CLis 0.3. While AOA=5, the smallest lift coefficient,CLis 0.15. While AOA=10, the smallest lift coefficient,CLis 0. According to this, while AOA=10, the lifting force,Lis no longer negative. Thus, the angle of attack moderate increasing will help the average lift coefficientCLincrease.In addition to the drag coefficient,CDin the downstroke and upstroke, the profile change of oscillation amplitude is obvious. When flapping wing starting downstroke and arriving the center point position from the highest peak, the drag coefficient,CDfalls to the lowest. Again wing flapping from the center point downstroke to the perigee position, the drag coefficient,CDelevates to the starting value, this may know while in downstroke the thrust force,Tis positive. Then wing flapping starts to upstroke from the perigee to the center point position, the drag coefficient,CDrises to the highest. The movement of wing flaps to upstroke again from the center point to the peak position, the drag coefficient,CDfalls to starting value, this means while in upstroke the thrust force,Tis also positive.Although in downstroke the minimum drag coefficient,CDassumes that the linear proportion to reduce, but it reduces relatively along with the angle of attack increase. While AOA=0, the minimum drag coefficient,CDis 0.018. While AOA=5, the minimum drag coefficient,CDis 0.06. While AOA=10, the minimum drag coefficient,CDis 0.135. But in upstroke the biggest drag coefficient,CDactually assumes that the linear proportional increasing. While AOA=0, the biggest drag coefficient,CDis 0.015. While AOA=5, the biggest drag coefficient,CDis 0.005. While AOA=10, the biggest drag coefficient,CDis 0.02. It increases along with the angle of attack increase, although in upstroke the biggest drag coefficient,CDdoes not assume that the linear proportion to reduce, but for all cases, the angle of attack increases will help the entire cyclical of the average thrust forceT.Wing tip vortexFigs. 7 and 9are the speed of vector diagrams for AOA=0 and AOA=10, whenK=0.3and at the position of 1/4 chord length observes the wing tip vortex. While AOA=0, regardless of in downstroke or upstroke, they