機(jī)械小黃鴨的機(jī)械機(jī)構(gòu)設(shè)計(jì)及運(yùn)動(dòng)仿真-機(jī)械鳥(niǎo)【Creo三維含10張CAD圖帶開(kāi)題報(bào)告-獨(dú)家】.zip
機(jī)械小黃鴨的機(jī)械機(jī)構(gòu)設(shè)計(jì)及運(yùn)動(dòng)仿真-機(jī)械鳥(niǎo)【Creo三維含10張CAD圖帶開(kāi)題報(bào)告-獨(dú)家】.zip,Creo三維含10張CAD圖帶開(kāi)題報(bào)告-獨(dú)家,機(jī)械,小黃鴨,機(jī)構(gòu),設(shè)計(jì),運(yùn)動(dòng),仿真,Creo,三維,10,CAD,開(kāi)題,報(bào)告,獨(dú)家
畢業(yè)論文附件材料
機(jī)械小黃鴨的機(jī)械機(jī)構(gòu)設(shè)計(jì)及運(yùn)動(dòng)仿真
學(xué)生姓名：
黃亞龍
學(xué) 號(hào)：
01410141Y31
所在系部：
機(jī)械工程系
專業(yè)班級(jí)：
14gb機(jī)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(xiàn)翻譯 1
1.1 Design and aerodynamic analysis of a flapping-wing micro aerial vehicle 1
Numerical analysis 4
1.2中文翻譯 25
2 專業(yè)閱讀書(shū)目 41
2.1 機(jī)械設(shè)計(jì) 41
2.2 機(jī)械原理 41
2.3 機(jī)電一體化系統(tǒng)設(shè)計(jì) 42
2.4 Creo2.0機(jī)械設(shè)計(jì)教程 42
2.5 單片機(jī)原理及接口技術(shù) 43
2.6 機(jī)械設(shè)計(jì)課程設(shè)計(jì) 43
2.7 現(xiàn)代工程制圖 44
2.8 機(jī)械零部件選用與設(shè)計(jì) 45
2.9 材料力學(xué) 46
2.10 理論力學(xué) 46
1 英文文獻(xiàn)翻譯
1.1 Design and aerodynamic analysis of a flapping-wing micro aerial vehicle
Author：Bor-JangTsai Yu-ChunFu
Abstract
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, Lˉ 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, Lˉ 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 Li–H 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 10–20 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 5–18 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 ornithopter's 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 Huang's [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. It's 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, Lˉ 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 analysis
Numerical method
In 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)
1g??t(gρ?)+div(ρu→r??Γ?grad?)=S?
ψ:On behalf of any independent physical quantity(ui,e,k…)
Γ?:Diffusion coefficient
S?:Source coefficient
After 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)C?k=∑(|BPn?Pn|?|BP0?P0|)<(given value)
Design budgetary estimates
Estimation of the MAV weight
The 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 estimates
Lift coefficient:
(3)CL=2LρU2S
Thrust coefficient:
(4)CT=2TρU2S
Reduced frequency:
(5)K=πfcU
Advance ratio:
(6)J=U2ΦfR=Flying speed or fluid velocityWing tip speed
Reynolds number:
(7)Re=Cˉ?Uˉtν=4ΦfR2ν?AR
Discussion and analysis of the numerical results.
Numerical simulation Geometry contour and grid establishment
In order to conform to DARPA's 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 Φ 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 different?K?AOA=0°,?K=0.1, 0.2, and 0.3
Lifting force and thrust force
When the angle of attack is 0° and the?K?value is 0.1, 0.2 and 0.3 respectively, investigates the increasing of?K?to influence on the aerodynamic forces.?Fig. 4?shows the comparison of lift coefficient,?CL?and different drag coefficient,?CD?based on different?K?values, 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,?CL?elevates to the maximum value, the movement of flap wing flapping again from the center point downstroke to the perigee position, and the lift coefficient,?CL?falls to the starting value. Therefore, in downstroke for the lifting force,?L?is positive. Starting to upstroke, the flap flapping from the perigee to the center point position, the lift coefficient,?CL?falls to the minimum value, the movement of flap wing flapping again from the center point to the peak position, the lift coefficient,?CL?rises to the starting value, thus the lifting force,?L?is negative value in upstroke.
The increase of?K?causes the profile of top and bottom oscillation amplitude for the lift coefficient,?CL?to become the proportional increasing, while in downstroke, the positive lift coefficient,?CL?becomes the proportion to increase. While?K=0.1, the maximum of lift coefficient,?CLis 0.1. While?K=0.2, the maximum of lift coefficient,?CL?is 0.2. While?K=0.3, the maximum of lift coefficient,?CL?is 0.3. While in upstroke, the negative lift
coefficient,?CL?becomes the proportional increasing. While?K=0.1, the smallest lift coefficient,?CL?is ?0.1. While?K=0.2, the smallest lift coefficient,?CL?is ?0.2. While?K=0.3, the smallest lift coefficient,?CL?is ?0.3. Increase of the positive and the negative counterbalances mutually, thus the?K?value 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,?CD?portion, while the?K?increases, the drag coefficient,?CD?has the big variation only when the flap starts flapping. While?K=0.1, the biggest drag coefficient,?CD?is ?0.0125. While?K=0.2, the biggest drag coefficient,?CD?is ?0.014. While?K=0.3, the biggest drag coefficient,?CD?is ?0.015. The drag coefficient,?CD?reduces relatively when the?K?value increases, after the first flap cycle, no matter how?K?value is, both in downstroke and in upstroke will not have big change, the mean drag coefficient,?CD?is ?0.018. As a result, while the angle of attack is 0°, the increase of?K?value does not have a quite big contribution to the average thrust coefficient.
Wing tip vortex
In 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=0/6?t/T=5/6 are in order.?Figs. 6 and 7?show the velocity vector diagrams for?K=0.1?and?K=0.3?respectively, at the position of 1/4 chord length observes the wing tip vortex. While the?t/T=0?starting 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,?L?for the plate wing. While the?t/T=3/6in the perigee position of downstroke, instantaneously, the turbulent flow can be absorbed because of the big reacting force. While the?t/T=4/6?starting 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,?L?is not favor for the MAV flight.
While?K=0.1, no matter how the downstroke or upstroke is, the wing tip vortex appears smooth. While?K=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 the?K?increase 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,?L?and thrust force,?T?as well. Regardless of the increasing of?K, the upstroke and downstroke have the same clockwise and counterclockwise strength of the vortex, therefore the positive and the negative of lifting force,?L?is mutually offset. This causes the average lifting force equal to zero. This result verifies that?CL?and?CD?of different?K?at 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 10°
Lifting force and thrust force
K=0.3, AOA?=?0°, 5° and 10°, investigates the increasing of?K?to influence on the lift coefficient,?CL?and the drag coefficient,?CD.?Fig. 8?is the comparison of the lift coefficient,?CL?and the drag coefficient,?CD?under the different angle of attack, so the increasing angle of attack conducive to favor the lifting force,?L?and the thrust force,?T?generation, while in downstroke the positive lift coefficient,?CL?becomes the proportion to increase. While AOA?=?0°, the maximum lift coefficient,?CLis 0.3. While AOA?=?5°, the maximum lift coefficient,?CL?is 0.5. While AOA?=?10°, the maximum lift coefficient,?CL?is 0.7. While in upstroke, the negative lift coefficient,?CL?becomes the proportional reducing actually. While the AOA?=?0°, the smallest lift coefficient,?CL?is ?0.3. While AOA?=?5°, the smallest lift
coefficient,?CL?is ?0.15. While AOA?=?10°, the smallest lift coefficient,?CL?is 0. According to this, while AOA?=?10°, the lifting force,?L?is no longer negative. Thus, the angle of attack moderate increasing will help the average lift coefficient?CˉL?increase.
In addition to the drag coefficient,?CD?in 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,?CD?falls to the lowest. Again wing flapping from the center point downstroke to the perigee position, the drag coefficient,?CD?elevates to the starting value, this may know while in downstroke the thrust force,?T?is 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,?CD?falls to starting value, this means while in upstroke the thrust force,?T?is also positive.
Although in downstroke the minimum drag coefficient,?CD?assumes that the linear proportion to reduce, but it reduces relatively along with
the angle of attack increase. While AOA?=?0°, the minimum drag coefficient,?CD?is ?0.018. While AOA?=?5°, the minimum drag coefficient,?CD?is ?0.06. While AOA?=?10°, the minimum drag coefficient,?CD?is ?0.135. But in upstroke the biggest drag coefficient,?CD?actually assumes that the linear proportional increasing. While AOA?=?0°, the biggest drag coefficient,?CD?is ?0.015. While AOA?=?5°, the biggest drag coefficient,?CD?is ?0.005. While AOA?=?10°, the biggest drag coefficient,?CD?is ?0.02. It increases along with the angle of attack increase, although in upstroke the biggest drag coefficient,?CD?does 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 force?Tˉ.
Wing tip vortex
Figs. 7 and 9?are the speed of vector diagrams for AOA?=?0° and AOA?=?10°, when?K=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
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