Design and optimization method of a high-lift rotor with low Reynolds number flow

Abstract

The present invention is a low Reynolds number flow high lift rotor design and optimization method. The present invention relates to the technical field of low Reynolds number flow high lift rotor design. The present invention selects a two-dimensional airfoil based on the profili airfoil database; determines the torsion angle and chord length based on the selected two-dimensional airfoil; determines the blade flow Reynolds number and selects the aspect according to the determined torsion angle and chord length; establishes a simulation model, selects a suitable turbulence model equation, and links the pulsation term and the time-averaged term to solve the equation; adopts the five-point method, continuously increases the number of grids for grid verification; modifies the model, and completes the design of a low Reynolds number flow high lift rotor. The nonlinear torsion design of the present invention has the best performance in improving lift, and its lift and unit volume lift are the highest among all models.

Description

Technical Field

[0001] The invention relates to the technical field of low Reynolds number flow large lift rotor design, and discloses a low Reynolds number flow large lift rotor design and optimization method. Background Art

[0002] Deep space exploration projects, mainly the Mars helicopter, have low flow Reynolds numbers due to the low pressure and low density of the aircraft's working environment. Conventional rotor designs find it difficult to obtain sufficient lift required for flight in such a rarefied environment. Therefore, improving the rotor lift under low Reynolds numbers has become a design focus.

[0003] Existing research results mostly focus on the low-pressure and low-density performance corrections of traditional rotors, but ignore the special atmospheric environment of Mars, namely the changes in flow sound speed caused by gas composition and day and night temperature changes, which will affect the reselection of two-dimensional airfoils and the determination of motor speed limits; at the same time, due to the small drag torque brought by low Reynolds number flow, the negative twist used in conventional rotors is no longer suitable for use due to poor lift performance. Summary of the invention

[0004] The present invention improves the obtained lift as much as possible by changing the rotor design means under the premise of considering the special flow environment of Mars. Therefore, the present invention provides a low Reynolds number flow large lift rotor design and optimization method. The present invention provides the following technical solutions:

[0005] A method for designing a low Reynolds number flow high lift rotor, the method specifically comprising:

[0006] Step 1: Select a two-dimensional airfoil based on the profili airfoil database;

[0007] Step 2: Based on the selected 2D airfoil, determine the twist angle and chord length;

[0008] Step 3: According to the determined twist angle and chord length, determine the blade flow Reynolds number and select the aspect chord;

[0009] Step 4: Establish a simulation model, select an appropriate turbulence model equation, and link the pulsation term and the time-averaged term to solve the equation;

[0010] Step 5: Use the five-point method to continuously increase the number of grids for grid verification;

[0011] Step 6: Modify the model and complete the design of the low Reynolds number flow high lift rotor.

[0012] Preferably, the step 1 is specifically:

[0013] Based on the profili airfoil database, a two-dimensional airfoil was selected, and the airfoil with a higher lift coefficient at a Reynolds number of 30,000 was selected. Then, a series of airfoils were selected for comparison based on geometric similarity, and the NASA SC 21010 airfoil with a higher lift coefficient at low Reynolds number and small torsion angle and the OAF095 airfoil with a higher lift coefficient at low Reynolds number and large torsion angle were selected.

[0014] Preferably, the step 2 is specifically:

[0015] Design of twist angle and chord length: The two selected airfoils each have their most suitable angle of attack. The lift coefficient of the NASA SC 21010 airfoil is greater when the angle of attack is less than 6°, and the lift coefficient of the OAF095 airfoil is greater when the angle of attack is greater than 6° and before the stall occurs at 11°. When the given twist angle increases with the increase of the rotation radius, that is, the positive twist law, a safety margin of 1° is reserved for each of the above two airfoils to prevent them from entering a stall state after being disturbed, causing a significant drop in lift.

[0016] The NASA SC 21010 airfoil is selected as the root of the blade with a twist angle of 0° to 5°; the OAF095 airfoil is selected as the tip of the blade with a twist angle of 5° to 10°. The Reynolds number of the segmented point flow is given as 15000 and the Reynolds number of the blade tip flow is given as 30000. The characteristic length of the flow at each point under a given Reynolds number, that is, the chord length of the airfoil and the distribution law of the Reynolds number extension length, are expressed by the following formula:

[0017]

[0018]

[0019] Preferably, the step 3 is specifically:

[0020] Selection of blade flow Reynolds number and aspect ratio. After determining the distribution law of blade twist angle and chord length extension direction, considering the overall size of the rotor, the blade aspect ratio is given as 10 according to the following formula (3), and the wingspan is determined to be 101.6 cm, that is, the span of a single blade is 50.8 cm;

[0021]

[0022] Wherein, L is the rotor wingspan, S is the top view area;

[0023] When the induced speed caused by the pressure difference generated by the rotation is not considered, the series rotor speed limit is about 4450rpm according to the following formula (4):

[0024]

[0025] Among them, n is the motor speed and r is the rotation radius.

[0026] Preferably, in the rotor flow field, when the rotation radius is small, the flow Reynolds number is also small, that is, laminar flow will occur, reducing the effective facing area of ​​the airfoil and thus affecting the lift. A horizontal axis with a single-side length of 37.5 mm is installed to avoid the laminar area.

[0027] Preferably, the step 4 is specifically:

[0028] Step 4.1: Determine a non-twisted NASA SC 21010 single-wing rotor, select the SA turbulence model, select an external static square domain with a size of 7200mm×5000mm×5000mm, set a horizontal flow with a speed of 10m / s on one side, and the remaining surfaces are free outflow boundary conditions. The static square computational domain mesh is generated by ICEM CFD;

[0029] Step 4.2: Use air as the working fluid, and the inlet flow velocity is 10m / s. At this time, the flow is simplified to a constant temperature incompressible motion, satisfying the control equation in fluid mechanics, the continuity equation:

[0030]

[0031] Momentum equation:

[0032]

[0033] Among them, u i is the velocity, i=1,2,3; p is the pressure; x i is the coordinate; f i is the mass force; t is the time; v is the kinematic viscosity;

[0034] The idea of ​​turbulent average motion is adopted, and the instantaneous value is expressed as the sum of the average value and the pulsation value, that is:

[0035]

[0036] Where D' is the pulsation value, is the average value, and D is the instantaneous value;

[0037] Conservation of mass:

[0038]

[0039] Conservation of Momentum:

[0040]

[0041] in: is the Reynolds stress;

[0042] Step 4.3: After the Reynolds stress is introduced, the number of equations is insufficient to solve; Select a suitable turbulence model equation, link the pulsation term and the time-averaged term to solve the equation, and use the SA model, a single-equation turbulence model suitable for airfoil and wall boundary layer flow and relatively simple to solve:

[0043]

[0044] Where v is the molecular viscosity; σ is the turbulent Prandtl number; Outside the transition layer and the viscous sublayer, it is equal to the eddy viscosity ratio v t ; for The corresponding strain rate tensor.

[0045] Preferably, the step 5 is specifically:

[0046] A torsion-free 1010 rotor was calculated using the five-point method, with the number of grids continuously increased. When the total number of grid cells was about 3×106, the calculation results were relatively accurate. Considering the computing resources, subsequent models all adopted this combination of 1×106 cells in the stationary domain and 2×106 cells in the rotating domain.

[0047] Preferably, the step 6 is specifically:

[0048] Considering the model correction of induced speed, the compensation is achieved by increasing the blade twist angle near the blade tip: it is approximately assumed that the speed V of the airflow at the blade tip before passing through the flow field is 20m / s, and V is 0 at the segmentation point. When the rotor angle of attack is kept unchanged, the second section of the blade needs to increase the twist angle correction of 5° at the blade tip, that is, the twist angle of the second section changes from 5° to 10° to 5° to 15°.

[0049] A nonlinear torsion design method for a low Reynolds number flow high lift rotor comprises the following steps:

[0050] The torsion angle of the rotor model changes linearly, that is, the torsion angle of each spanwise micro-segment is evenly distributed along the spanwise direction. However, the relationship between the lift coefficient and the torsion angle of each existing airfoil is not always linear. For the selected airfoil NASA SC 21010, the lift coefficient at an angle of attack of 3° to 5° is significantly higher than that at 0° to 3°; for the selected airfoil OAF095, its lift coefficient at an angle of attack of 7° to 10° is also higher than that at 5° to 7°;

[0051] For the two selected airfoils, since the lift coefficient is higher and the resulting lift is greater at larger twist angles, the area enclosed by the blue line and the x-axis is proportional to the lift obtained by the rotor. That is, under the premise of given functions and constraint points, a set of parameters that meet the maximum enclosed area requirement can be found.

[0052] Preferably, in order to minimize the calculation time, a selected function is given for each segment, and origin is used to roughly give the initial value of the function. The optimization is to find the maximum value near the initial value.

[0053] The present invention has the following beneficial effects:

[0054] The nonlinear torsion design of the present invention has the best performance in improving lift. Both its lift and unit volume lift are the highest among all models. Compared with the linear torsion arrangement, its unit volume lift is increased by 14.72%; the lift and unit volume lift of the center-aligned linear torsion design model are higher than those of the leading edge-aligned model, with gains of 3.17% and 3.21% respectively; the two-stage design has achieved better results in improving lift than the single-stage design, but the drag moment has also increased accordingly: NASA SC(2)1010 and OAF095 airfoils have obtained unit volume lift gains of 22.42% and 17.90% respectively, while the drag moment has increased by 14.03% and 7.24% respectively; compared with the single airfoil, the linear design model composed of the two airfoils has achieved a better compromise performance: its unit volume lift is slightly lower than that of NASA The SC(2)1010 airfoil (decreased by 6.42%), but at the same time reduced the drag moment by 14.23%; compared with the OAF095 airfoil, its drag moment increased by 17.8%, but the lift increased by 15.68%; due to the obvious irrationality of the non-twisted arrangement, its lift and unit volume lift are much lower than those of the aforementioned models.

[0055] When the change in angle of attack caused by the induced speed is taken into account, the lift per unit volume of the linear torsion model increases by 20.92%; compared with the nonlinear torsion arrangement, the lift per unit volume increases by 5.43%, while the drag moment increases by only 2.01%, which also shows that for the rotor flow field with low Reynolds number flow, the arrangement in which the torsion angle extends to increase, that is, positive torsion, is reasonable. For the negative torsion arrangement currently used by most helicopters, since its large torsion angle blade element works at the root of the blade with a lower Reynolds number, and the negative torsion change trend at the blade tip will further reduce the effective angle of attack, resulting in very low lift and unit volume lift. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] Figure 1 Select a schematic diagram for the airfoil;

[0057] Figure 2 Design diagram for the torsion angle and chord length of the blade;

[0058] Figure 3 is the lift coefficient diagram;

[0059] Figure 4 To design the overall picture of the model;

[0060] Figure 5 To compare the static computational domain diagrams of four different heights;

[0061] Figure 6 The calculation results of SA turbulence model are selected;

[0062] Figure 7 is the grid cell data;

[0063] Figure 8 The strength of the induction effect;

[0064] Fig. 9 This is a schematic diagram of nonlinear torsion design;

[0065] Fig.10 More blade elements work at a larger lift coefficient;

[0066] Fig.11 is the distribution diagram of the torsion angle and the rotation radius of nonlinear torsion;

[0067] Fig.12 Take screenshots of your work;

[0068] Fig.13 Take screenshots of your work;

[0069] Fig.14 Take screenshots of your work;

[0070] Fig.15 Take screenshots of your work;

[0071] Fig.16 Take screenshots of your work;

[0072] Fig.17 Take screenshots of your work;

[0073] Fig.18 Find such a maximum value near the initial value;

[0074] Fig.19 The nonlinear torsion design has the best performance in terms of lift improvement;

[0075] Fig. 20 is the lift per unit volume of the linear torsion model when the change in angle of attack caused by the induced velocity is taken into account. DETAILED DESCRIPTION

[0076] The present invention is described in detail below in conjunction with specific embodiments. Specific embodiment one:

[0078] according to Figures 1 to 20 As shown, the specific optimization technical solution adopted by the present invention to solve the above technical problems is: the present invention relates to a low Reynolds number flow high lift rotor design and optimization method.

[0079] A method for designing a low Reynolds number flow high lift rotor, the method specifically comprising:

[0080] Step 1: Select a two-dimensional airfoil based on the profili airfoil database;

[0081] The step 1 is specifically as follows: based on the profili airfoil database, a two-dimensional airfoil is selected, an airfoil with a higher lift coefficient at a Reynolds number of 30,000 is selected, and then a series of airfoils are selected for comparison based on geometric similarity, and the NASA SC(2)1010 airfoil with a higher lift coefficient at a low Reynolds number and a small twist angle and the OAF095 airfoil with a higher lift coefficient at a low Reynolds number and a large twist angle are selected. The selection of airfoils is shown in Figure 1 .

[0082] Step 2: Based on the selected 2D airfoil, determine the twist angle and chord length;

[0083] The step 2 is specifically as follows:

[0084] Design of twist angle and chord length: The two selected airfoils each have their most suitable angle of attack. The lift coefficient of the NASA SC(2)1010 airfoil is greater when the angle of attack is less than 6°, and the lift coefficient of the OAF095 airfoil is greater when the angle of attack is greater than 6° and before the stall occurs at 11°. When the given twist angle increases with the increase of the rotation radius, that is, the positive twist law, a safety margin of 1° is reserved for each of the above two airfoils to prevent them from entering a stall state after being disturbed, causing a significant drop in lift.

[0085] The NASA SC(2)1010 airfoil is selected as the root of the blade with a twist angle of 0° to 5°; the OAF095 airfoil is selected as the tip of the blade with a twist angle of 5° to 10°. The Reynolds number of the segment flow is given as 15000 and the Reynolds number of the blade tip flow is given as 30000. The characteristic length of the flow at each point under a given Reynolds number, that is, the chord length of the airfoil and the distribution law of the Reynolds number extension length, are obtained, which are expressed by the following formula:

[0086]

[0087] Step 3: According to the determined twist angle and chord length, determine the blade flow Reynolds number and select the aspect chord;

[0088] The step 3 is specifically as follows:

[0089] Selection of blade flow Reynolds number and aspect ratio. After determining the distribution law of blade twist angle and chord length extension direction, considering the overall size of the rotor, the blade aspect ratio is given as 10 according to the following formula (3), and the wingspan is determined to be 101.6 cm, that is, the span of a single blade is 50.8 cm;

[0090]

[0091] Wherein, L is the rotor wingspan, S is the top view area;

[0092] When the induced speed caused by the pressure difference generated by the rotation is not considered, the series rotor speed limit is about 4450rpm according to the following formula (4):

[0093]

[0094] Among them, n is the motor speed and r is the rotation radius.

[0095] In the rotor flow field, when the rotation radius is small, the flow Reynolds number is also small, that is, laminar flow will occur, reducing the effective positive area of ​​the airfoil and thus affecting the lift. A horizontal axis with a single-side length of 37.5 mm is added to avoid the laminar area. The overall design model is shown in Figure 4 .

[0096] Step 4: Establish a simulation model, select an appropriate turbulence model equation, and link the pulsation term and the time-averaged term to solve the equation;

[0097] The step 4 is specifically as follows:

[0098] Step 4.1: Considering the influence of the horizontal wind on Mars and the axial flow of the rotor flow field, the static calculation domains at four different heights are compared. Figure 5 , a non-twisted NASA SC 21010 single-wing rotor is determined, and the SA turbulence model is selected. The calculation results are as follows Figure 6 An external static square domain with a size of 7200mm×5000mm×5000mm was selected, and a horizontal flow with a velocity of 10m / s was set on one side. The remaining surfaces were all free outflow boundary conditions. The static square computational domain mesh was generated by ICEM CFD.

[0099] Step 4.2: Use air as the working fluid, and the inlet flow velocity is 10m / s. At this time, the flow is simplified to a constant temperature incompressible motion, satisfying the control equation in fluid mechanics, the continuity equation:

[0100]

[0101] Momentum equation:

[0102]

[0103] Among them, u i is the velocity, i=1,2,3; p is the pressure; x i is the coordinate; f i is the mass force; t is the time; v is the kinematic viscosity;

[0104] The idea of ​​turbulent average motion is adopted, and the instantaneous value is expressed as the sum of the average value and the pulsation value, that is:

[0105]

[0106] Where D' is the pulsation value, is the average value, and D is the instantaneous value;

[0107] Conservation of mass:

[0108]

[0109] Conservation of Momentum:

[0110]

[0111] in: is the Reynolds stress;

[0112] Step 4.3: After the Reynolds stress is introduced, the number of equations is insufficient to solve; Select a suitable turbulence model equation, link the pulsation term and the time-averaged term to solve the equation, and use the SA model, a single-equation turbulence model suitable for airfoil and wall boundary layer flow and relatively simple to solve:

[0113]

[0114] Where v is the molecular viscosity; σ is the turbulent Prandtl number; Outside the transition layer and the viscous sublayer, it is equal to the eddy viscosity ratio v t ; for The corresponding strain rate tensor.

[0115] Step 5: Use the five-point method to continuously increase the number of grids for grid verification;

[0116] The step 5 is specifically as follows:

[0117] The same as the above selection of static calculation domain, a non-twisted 1010 rotor is calculated and the five-point method is used to continuously increase the number of grids. The grid unit data of each part are shown in Table 1. The results are shown in Figure 7 When the total number of grid cells is about 3×106, the calculation results are relatively accurate. Considering the computing resources, the subsequent models all use this combination of 1×106 cells in the stationary domain and 2×106 cells in the rotating domain.

[0118] Table 1. Grid cell data

[0119]

[0120] Step 6: Modify the model and complete the design of the low Reynolds number flow high lift rotor.

[0121] The step 6 is specifically as follows:

[0122] Model correction considering induced speed: To obtain lift, the rotor must discharge part of the airflow downward. This effect of the rotor disc on the airflow is called "induction". The velocity increment obtained after the airflow passes through the rotor flow field along the axial direction is called "induced speed". Figure 8 As can be seen in the lower right corner, the strength of this induced effect is not uniformly distributed in the extension direction: even in the hovering state, the initial velocity of the airflow passing through the propeller disk from top to bottom is not always 0, and the effect is more obvious the closer to the propeller tip. At this time, the change in the angle of attack caused by the airflow blowing along the axis to the blade cannot be ignored.

[0123] The compensation is achieved by increasing the blade twist angle near the blade tip: it is approximately assumed that the speed V of the airflow at the blade tip before passing through the flow field is 20m / s, and V is 0 at the segmentation point. When the rotor angle of attack is kept unchanged, the second section of the blade needs to increase the twist angle correction by 5° at the blade tip, that is, the twist angle of the second section changes from 5° to 10° to 5° to 15°.

[0124] A nonlinear torsion design method for a low Reynolds number flow high lift rotor comprises the following steps:

[0125] The torsion angle of the rotor model changes linearly, that is, the torsion angle of each spanwise micro-segment is evenly distributed along the spanwise direction. However, the relationship between the lift coefficient and the torsion angle of each existing airfoil is not always linear. For the selected airfoil NASA SC 21010, the lift coefficient at an angle of attack of 3° to 5° is significantly higher than that at 0° to 3°; for the selected airfoil OAF095, its lift coefficient at an angle of attack of 7° to 10° is also higher than that at 5° to 7°;

[0126] For the two selected airfoils, since the lift coefficient is higher and the resulting lift is greater at larger twist angles, the area enclosed by the blue line and the x-axis is proportional to the lift obtained by the rotor. That is, under the premise of given functions and constraint points, a set of parameters that meet the maximum enclosed area requirement can be found.

[0127] In order to minimize the calculation time, the selected function for each segment is given, and the origin is used to roughly give the initial value of the function. The optimization is to find the maximum value near the initial value.

[0128] Nonlinear torsion design method: The torsion angle of the above-mentioned rotor model changes linearly, that is, the torsion angle of each spanwise micro-segment (the so-called "blade element") is evenly distributed along the spanwise direction. However, the relationship between the lift coefficient and the torsion angle of each existing airfoil is not always linear. For the selected airfoil NASA SC(2)1010, its lift coefficient at an angle of attack of 3° to 5° is significantly higher than that at 0° to 3°; for the selected airfoil OAF095, its lift coefficient at an angle of attack of 7° to 10° is also higher than that at 5° to 7°. On the lift coefficient diagram, 3° and 7° are called "large lift coefficient lines", so the torsion angle change curve of the blade is made to quickly pass through these two lines and transition as smoothly as possible at the segmentation point (such as Fig.10 ), more blade elements will work under a larger lift coefficient.

[0129] From the perspective of improving lift, this nonlinear torsion design is obviously more reasonable. This paper names this design method as the "fast crossing method". A nonlinear torsion model designed based on this method is as follows: Fig.11 .

[0130] Nonlinear torsion optimization procedure: The distribution of the torsion angle along the rotation radius of nonlinear torsion is as follows: Fig.12 For the two selected airfoils, since the lift coefficient is higher and the resulting lift is greater at a large twist angle, we can get Fig.12 The area enclosed by the blue line and the x-axis is proportional to the lift obtained by the rotor, that is, given the function and the constraint points, a set of parameters that satisfies the maximum enclosed area can be found. Such an optimization program was written based on Matlab, and its working screenshot is shown in Figures 12 to 17 .

[0131] In order to minimize the calculation time, the optimization program needs to first give the selected function of each segment and use origin to give a rough initial value of the function. The optimization is to find such a maximum value near the initial value. The result is shown in Fig.18 .

[0132] In summary, starting from the special low Reynolds number flow on Mars, the present invention proposes a two-dimensional low Reynolds number airfoil optimization method based on the profili airfoil platform and a three-dimensional linear / nonlinear blade design method, and completes the writing of a function parameter optimization program in nonlinear torsion.

[0133] Depend on Fig.19The nonlinear torsion design has the best performance in terms of lift improvement. Its lift and unit volume lift are the highest among all models. Compared with the linear torsion arrangement, its unit volume lift is increased by 14.72%. The lift and unit volume lift of the center-aligned linear torsion design model are higher than those of the leading edge-aligned model, with gains of 3.17% and 3.21% respectively. The two-stage design has achieved better results in lift improvement than the single-stage design, but the drag moment has also increased accordingly: NASA SC(2)1010 and OAF095 airfoils have gained 22.42% and 17.90% unit volume lift gains respectively, while the drag moment has increased by 14.03% and 7.24% respectively. Compared with the single airfoil, the linear design model composed of the two airfoils has achieved a better compromise performance: its unit volume lift is slightly lower than that of NASA The SC(2)1010 airfoil (decreased by 6.42%), but at the same time reduced the drag moment by 14.23%; compared with the OAF095 airfoil, its drag moment increased by 17.8%, but the lift increased by 15.68%; due to the obvious irrationality of the non-twisted arrangement, its lift and unit volume lift are much lower than those of the aforementioned models.

[0134] Depend on Fig. 20 , when the angle of attack changes caused by the induced speed are taken into account, the lift per unit volume of the linear torsion model increases by 20.92%; compared with the nonlinear torsion arrangement, the lift per unit volume increases by 5.43%, while the drag moment increases by only 2.01%, which also shows that for the rotor flow field with low Reynolds number flow, the arrangement in which the torsion angle extends to increase, that is, positive torsion, is reasonable. For the negative torsion arrangement currently used by most helicopters, since its large torsion angle blade element works at the root of the blade with a lower Reynolds number, and the negative torsion change trend at the blade tip will further reduce the effective angle of attack, resulting in very low lift and unit volume lift.

[0135] The above is only a preferred implementation of a low Reynolds number flow high lift rotor design and optimization method. The protection scope of a low Reynolds number flow high lift rotor design and optimization method is not limited to the above embodiments. All technical solutions under this idea belong to the protection scope of the present invention. It should be pointed out that for those skilled in the art, several improvements and changes without departing from the principle of the present invention should also be regarded as the protection scope of the present invention.

Claims

1. A method for designing a low Reynolds number flow high lift rotor, characterized by: The method is specifically as follows: Step 1: Select a two-dimensional airfoil based on the profili airfoil database; Step 2: Based on the selected 2D airfoil, determine the twist angle and chord length; The step 2 is specifically as follows: The NASA SC 21010 airfoil is selected as the root of the blade with a twist angle of 0° to 5°; the OAF095 airfoil is selected as the tip of the blade with a twist angle of 5° to 10°. The Reynolds number of the segmented point flow is given as 15000 and the Reynolds number of the blade tip flow is given as 30000. The characteristic length of the flow at each point under a given Reynolds number, that is, the chord length of the airfoil and the distribution law of the Reynolds number extension length, are expressed by the following formula: Step 3: According to the determined twist angle and chord length, determine the blade flow Reynolds number and select the aspect chord; The step 3 is specifically as follows: Selection of blade flow Reynolds number and aspect ratio. After determining the distribution law of blade twist angle and chord length extension direction, considering the overall size of the rotor, the blade aspect ratio is given as 10 according to the following formula (3), and the wingspan is determined to be 101.6 cm, that is, the span of a single blade is 50.8 cm; Wherein, L is the rotor wingspan, S is the top view area; When the induced speed caused by the pressure difference generated by the rotation is not considered, the series rotor speed is limited to 4450rpm according to the following formula (4): Where n is the motor speed and r is the rotation radius; Step 4: Establish a simulation model, select the turbulence model equation, and link the pulsation term and the time-averaged term to solve the equation; The step 4 is specifically as follows: Step 4.1: Determine a non-twisted NASA SC 21010 single-wing rotor, select the SA turbulence model, select an external static square domain with a size of 7200mm×5000mm×5000mm, set a horizontal flow with a speed of 10m / s on one side, and the remaining surfaces are free outflow boundary conditions. The static square computational domain mesh is generated by ICEM CFD; Step 4.2: Use air as the working fluid, and the inlet flow velocity is 10m / s. At this time, the flow is simplified to a constant temperature incompressible motion, satisfying the control equation in fluid mechanics, the continuity equation: Momentum equation: Among them, u i is the velocity, i=1,2,3; p is the pressure; x i is the coordinate; f i is the mass force; t is the time; v is the kinematic viscosity; The idea of ​​turbulent average motion is adopted, and the instantaneous value is expressed as the sum of the average value and the pulsation value, that is: Where D' is the pulsation value, is the average value, and D is the instantaneous value; Conservation of mass: Conservation of Momentum: in: is the Reynolds stress; Step 4.3: After introducing Reynolds stress, select the single equation turbulence model SA model: Where v is the molecular viscosity; σ is the turbulent Prandtl number; Outside the transition layer and the viscous sublayer, it is equal to the eddy viscosity ratio v t ; for The corresponding strain rate tensor; Step 5: Use the five-point method to continuously increase the number of grids for grid verification; Step 6: Modify the model and complete the design of the low Reynolds number flow high lift rotor; The step 6 is specifically as follows: Set the velocity V of the airflow at the blade tip before it passes through the flow field to 20m / s and V at the segmentation point to 0. While keeping the rotor angle of attack unchanged, the second section of the blade needs to increase the torsion angle correction by 5° at the blade tip, that is, the torsion angle of the second section changes from 5° to 10° to 5° to 15°.

2. A method for designing a low Reynolds number flow high lift rotor according to claim 1, characterized in that: The step 1 is specifically as follows: Based on the profili airfoil database, a two-dimensional airfoil was selected. The airfoil with a high lift coefficient at a Reynolds number of 30,000 was selected. Then, a series of airfoils were selected for comparison based on geometric similarity. The NASA SC 21010 airfoil with a high lift coefficient at a low Reynolds number and a small torsion angle and the OAF095 airfoil with a high lift coefficient at a low Reynolds number and a large torsion angle were selected.

3. The method for designing a low Reynolds number flow high lift rotor according to claim 1, characterized in that: In the rotor flow field, a horizontal axis with a single-side length of 37.5 mm was installed to avoid the laminar flow area.

4. The method for designing a low Reynolds number flow high lift rotor according to claim 1, characterized in that: The step 5 is specifically as follows: A torsion-free 1010 rotor is calculated and the five-point method is used to increase the number of grids. When the total number of grid elements is 3×106, subsequent models all use this combination of 1×106 elements in the stationary domain and 2×106 elements in the rotating domain.

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