CN115983014A - Design method for single-rotating-shaft lift margin aircraft adjustable wing based on geometric strong constraint - Google Patents
Design method for single-rotating-shaft lift margin aircraft adjustable wing based on geometric strong constraint Download PDFInfo
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Abstract
The invention discloses a design method of an adjustable wing of a single-rotating-shaft lift margin aircraft based on geometric strong constraint, which comprises the following three steps: determining the position range of a rotation center, the design flow of a sectional type adjustable wing and the checking of the maximum stress area of a deformation wing; determining the position range of the rotating center by combining the input aircraft model, the internal load and the radius of the rotating shaft, taking points at intervals above the rotating center, and designing adjustable wings by taking each point as the center of the rotating shaft; the position range is divided into two or three areas, and different design flows are adopted for points in different areas; and selecting a scheme capable of maximizing the unfolding area of the wing as a design result, calculating the stress of the wing under the condition of the extreme flight of the aircraft, checking the stress intensity to determine the minimum thickness of the wing, and finishing the design of the adjustable wing. The invention designs the deformation wing which rotates around the single shaft and has the largest area within the corresponding size envelope range based on different shapes of the external shape and the internal load of the aircraft.
Description
Technical Field
The invention relates to the technical field of hypersonic variant aircrafts, in particular to a design method of an adjustable wing of a single-rotating-shaft lift margin aircraft based on geometric strong constraint.
Background
The hypersonic variant aircraft is an aircraft with the flying speed of more than 5Ma and the shape variable, and has the characteristics of high-speed flight and excellent flight environment adaptability. With the development of aerospace technology and the diversification of flight requirements of various field departments, the requirement on aircraft performance is no longer limited to a specific flight altitude and mach number, i.e., the capability of large airspace and wide-speed-range flight, which makes the aircraft one of the research hotspots in the aerospace field rapidly.
In the current research on hypersonic aircrafts in large airspace and wide speed area at home and abroad, in order to ensure the low resistance characteristic and the lift-drag ratio, the appearances of the aircrafts are mostly characterized by large slenderness ratio, so that the volume ratio of the aircrafts is not very high. In practice, aircraft interiors are loaded with interior loads of a size that further compresses the otherwise available space within the elongated aircraft interior. How to design a variable wing with a larger area by using the limited volume becomes a key problem to be faced in the practical application of the hypersonic speed variant aircraft.
While the current research on variable wings is mainly focused on the impact on the overall aerodynamic performance of the aircraft, on the research of new deformation solutions, on the wing shape with wide lift-to-drag ratio margins and on the reliability of the deformation mechanisms, the use of the internal volume is also considered relatively seldom.
Disclosure of Invention
The invention aims to provide a design method of a single-rotating-shaft type high-speed aircraft adjustable wing with wide lift margin based on geometric strong constraint, which is used for corresponding to a deformed wing which rotates around a single shaft and has the largest area within a size envelope range under the strong constraint condition of given any body shape and internal loading, so that the lift-drag ratio of the aircraft when the wings are completely unfolded and fly and the lift margin which can be reached by the aircraft as a whole are improved.
In order to achieve the purpose, the invention provides the following technical scheme: the design method of the single-rotating-shaft lift margin aircraft adjustable wing based on the geometric strong constraint comprises the following steps:
s1, determining a position range of a rotary shaft center according to input data, and taking a series of points as the rotary shaft center;
s2, dividing the position range into two to three areas, and obtaining corresponding design results by adopting different design processes for points in different areas;
and S3, selecting the one which can maximize the expansion area value from the design results of the different points as a final design result, calculating the stress of the wing according to the extreme flight condition of the aircraft, checking the stress intensity to determine the minimum thickness h which the wing should take, and finishing the design of the adjustable wing.
Preferably, the location range and the point set { O } in step 1 0 ～O n The determination method is as follows:
1) Firstly, half-cutting an aircraft along the plane of a wing, taking an input radius r as an inner equal-distance line L of an outer contour line of an aircraft body, wherein the center of a rotating shaft is positioned on the L, and for the input radius r of the rotating shaft, a direct gap between the outer contour line of the aircraft body and an internal load is used for accommodating the rotating shaft;
2) On the curve L, n +1 points at certain intervals are taken as the center of the rotating shaft and are marked as { O 0 ～O n In which the starting position O 0 R from the central axis, i.e. z = r; end position O n Should be taken away from the bottom surface r of the circular truncated cone, namely x = x load1 -r。
Preferably, the location range division and segment design process in step 2 is as follows:
1) Dividing L into two or three sections according to the geometrical shape of the internal load;
2) Design of the deformed wing according to a two-stage/three-stage method, thereby determining O at each point i The corresponding wing rotating shaft, the outer contour of the wing,Bottom contour, initial wing inner contour, and O-winding the wing shape obtained by the above steps i Rotating anticlockwise to obtain the result when the wing is completely unfolded and recording the completely unfolded area S i And the inner profile of the wing is always positioned in the machine body in the rotating process, and the sectional type specifically comprises the following three steps:
①O i at L 3 Designing a flow of adjustable wings;
②O i at L 2 Designing a flow of adjustable wings;
③O i at L 1 The design process of the adjustable wing.
Preferably, the stress analysis and stress check of the wing in step 3 are as follows:
(1) Selecting a maximum expansion area S k ＝{S 0 ～S n Calculating the pressure difference (P) between the upper surface and the lower surface of the wing according to the area and the limit flight condition of the aircraft 3 -P 2 )；
(2) Uniformly distributing pressure (P) according to normal direction of wing 3 -P 2 ) And the wing is made of a material to be used, and the stress of the wing is checked, so that the minimum thickness h of the wing is determined to meet the following conditions:
wherein l is the maximum length of the outer contour of the wing from the outer contour line of the body, and [ sigma ] is the allowable stress of the wing material.
Compared with the prior art, the invention has the beneficial effects that:
1. the design method of the single-rotating-shaft type high-speed aircraft adjustable wing with wide lift margin based on the geometric strong constraint provided by the invention is used for designing the adjustable wing based on the shape of any given aircraft, the shape of internal loading and the position of an end point, so that the maximum area of the single-rotating-shaft type high-speed aircraft deformable wing in the corresponding size envelope range is designed under the geometric strong constraint.
2. The design method is suitable for the design of the single-rotating-shaft adjustable wing under any aircraft shape; the inner part is loaded with a circular truncated cone and a cylinder, and most loading forms in practical application are covered; the setting of the end point avoids the situation that the designed adjustable wing interferes with flying parts possibly carried by aircrafts such as air rudders and the like in the unfolding process.
3. After the design is finished and the requirements of practical application are considered, the minimum thickness required to be adopted by the adjustable wing can be roughly determined by performing stress analysis and stress check on the adjustable wing in the extreme flight state, and the application reliability of the adjustable wing is improved.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the principles of the invention and not to limit the invention.
In the drawings:
FIG. 1 is a three-dimensional schematic view of a hypersonic aircraft suitable for use with the present invention;
FIG. 2 is a general design flow diagram of the present invention;
FIG. 3 is a flow chart of a segmented design according to the present invention;
FIG. 4 is a schematic view of the internal loading shape and size for which the present invention is applicable;
FIG. 5 is a half sectional view of the aircraft in the step of determining the range of the center position of the adjustable wing rotor shaft according to the present invention;
FIG. 6 is a schematic diagram of the determination of the position range in the step of determining the position range of the center of the rotating shaft of the adjustable wing according to the present invention, i.e. the center of the rotating shaft needs to be located on the equidistant line L of the outer contour line of the aircraft body;
FIG. 7 is a diagram illustrating the determination of the position range in the step of determining the position range of the center of the adjustable wing shaft according to the present invention, based on the diagram of FIG. 6, the center of the adjustable wing shaft should be taken at L, O 0 To O n Schematic diagram of (a);
FIG. 8 is a three-step division diagram of the position range L in the step of determining the central position range of the rotating shaft of the adjustable wing according to the present invention;
FIG. 9 is a two-step division view of the position range L in the step of determining the range of the center position of the rotating shaft of the adjustable wing according to the present invention;
FIG. 10 shows the sectional design of the present invention with the center of the rotating shaft at L 3 A schematic diagram of a generation method of the upper time adjustable wing;
FIG. 11 shows the sectional design of the present invention with the center of the rotating shaft at L 3 Generating a result graph of the upper time adjustable wing;
FIG. 12 shows the sectional design of the present invention with the center of the rotating shaft at L 3 The maximum rotation angle of the adjustable wing during mounting;
FIG. 13 shows the sectional design of the present invention with the center of the rotating shaft at L 3 Schematic view of the adjustable wing when up fully deployed;
FIG. 14 shows the sectional design of the present invention with the center of the rotating shaft at L 2 A schematic diagram of a generation method of the upper time adjustable wing;
FIG. 15 shows the present invention in a segmented design flow with the center of the shaft at L 2 Generating a result graph of the upper time adjustable wing;
FIG. 16 shows the sectional design flow of the present invention in which the center of the rotating shaft is located at L 2 The maximum rotation angle of the adjustable wing during mounting;
FIG. 17 shows the present invention in a segmented design flow with the center of the shaft at L 2 Schematic view of the adjustable wing when up fully deployed;
FIG. 18 shows the present invention in a segmented design flow with the shaft centered at L 1 A schematic diagram of a generation method of the upper time adjustable wing;
FIG. 19 shows the present invention in a segmented design flow with the shaft centered at L 1 Generating a result graph of the upper time adjustable wing;
FIG. 20 shows the sectional design of the present invention with the center of the rotating shaft at L 1 The maximum rotation angle of the adjustable wing during mounting;
FIG. 21 shows the sectional design of the present invention with the center of the rotating shaft at L 1 Schematic view of the adjustable wing when up fully deployed;
FIG. 22 is a schematic view of a flow field for aerodynamic analysis of an adjustable wing in stress analysis and stress verification according to the present invention;
FIG. 23 is a schematic diagram of finding the maximum stress position of the adjustable wing during stress analysis and stress check according to the present invention;
FIG. 24 is a schematic view of the present invention during force analysis and stress verification for performing mechanical analysis on the adjustable wing to verify the bending stress;
FIG. 25 is a schematic diagram of the present invention during force analysis and stress verification to verify the bending stress of the adjustable wing;
FIG. 26 is a half sectional view in the xOz plane of the aircraft of FIG. 1 with the origin at the apex in accordance with an embodiment of the present invention;
FIG. 27 is a schematic view showing a range of positions where the center of the rotating shaft is desirable in determining the center of the rotating shaft in the embodiment of the present invention;
FIG. 28 is a graph of design results obtained with an embodiment of the present invention;
reference numbers in the figures: 1. an aircraft body; 2. an adjustable wing; 3. carrying out internal loading; 4. an air vane.
Detailed Description
The preferred embodiments of the present invention will be described in conjunction with the accompanying drawings, and it will be understood that they are described herein for the purpose of illustration and explanation and not limitation.
As shown in fig. 1, which is a model of an embodiment of the present invention, the wing of which is the design result of the embodiment of the present invention, for the high-speed aircraft with internal loading as shown in fig. 1, the internal loading shape can be mostly classified into 5 types as shown in fig. 4: cylinder, circular cone + cylinder, round platform and cylinder + round platform. The first 4 types of the wing-adjustable wing-shaped structure can be regarded as special forms of a cylinder and a circular truncated cone, and the design of the adjustable wing is carried out in the internal loading form of the cylinder and the circular truncated cone.
Referring to fig. 2, the design method of the single-rotating-shaft wide-lift-margin high-speed aircraft adjustable wing based on the geometric strong constraint comprises the following steps:
1) Referring to fig. 5, the body is half-sectioned at xOz facing the body with the head of the body as the origin.
Referring to fig. 6, the input radius r is taken as the inner equidistant line L of the outer contour line of the body, and the center of the rotating shaft is located on L. And the radius r of the input rotating shaft is to enable the gap between the outer contour line of the machine body and the inner load to accommodate the rotating shaft. And secondly, in order to prevent the designed adjustable wing from interfering with the air rudder in the rotation process, the vertex of the air rudder is used as an end point N of the adjustable wing.
2) Referring to FIG. 7, n +1 points are marked as { O ] at intervals on the curve L 0 ～O n }. Wherein the starting position O 0 R from the central axis, i.e. z = r; end position O n Should be taken away from the bottom surface r of the circular table, i.e. x = x load1 -r. For O optionally taken therein i ∈{O 0 ～O n Make a circle C with radius r i I.e. the rotating shaft.
1) L needs to be segmented according to the shape characteristics of the internal loading, which is specifically divided into two cases: when the extensions of AB are not all located at C 0 In the lower part, L should be divided into three sections, otherwise two sections.
Referring to FIG. 8, the rotation center O is taken at L 1 When the rotating shaft is integrally positioned at the straight line z = z load1 Upper side, L 1 Indicated in the figure by thick solid lines; the centre of rotation O being taken at L 2 When the rotating shaft is wholly positioned above the AB extension line, L 2 Indicated in the figure by a dash-dot line; the remaining curves up to O 0 Is represented by L 3 Indicated by a thin solid line.
Referring to fig. 9, the two-step method only needs to be based on the straight line z = z load1 To perform the division.
2) Referring to FIG. 3, it corresponds to the process of FIG. 2 where "the center of the rotation axis O is located at L 3 In the design method of (1), the center O of the rotating shaft is located at L 2 The design method of (1) 'and' the center O of the rotating shaft is located at L 1 The design method of (1), after the division is completed, the rootAccording to O i The design of the deformation wing is carried out at the position, and O is considered in the three-section type division method i Are respectively located at L 3 、L 2 、L 1 The above three cases, namely, the following steps (1), (2) and (3); while the two-stage division method only considers O i At L 2 、L 1 The above two cases, that is, only the following (2) and (3) need to be performed.
①O i At L 3 The method comprises the following steps:
referring to FIG. 10, L 3 Last point O i The distance from the point B of the internal load is recorded as a, and a tangent circle C with the point as the center of the circle and the outer contour line z = f (x) of the machine body is formed i . Then O i As the circle center, the radius from the center to the end point N is taken as an arc D, and at the moment, the straight line z = z above the internal load load1 The intersection point with the arc D is marked as N'. Firstly, O is added i And an inner loading vertex A (x) load0 ,z load0 ) Are connected to form 1 Connecting the point A and the point B to form a point I 2 Continue to connect vertex B (x) load1 ,z load1 ) And N' is l 3 。
Referring to FIG. 11, at this time, the shape of the airfoil is initially determined as a straight line l 1 Straight line l 2 Straight line l 3 Circular arc
Body outer contour line z = f (x), and circle C i And the enclosed black area. Namely, the wing rotating shaft, the wing outer contour and the wing bottom contour are determined.
Referring to fig. 12, in order to determine the maximum rotation angle of the adjustable wing, all the points, i.e., a, B, and N', which may be in initial contact with the outer contour line of the body are considered. First pass through N' and C i Tangent line l 5 Then go through the end points N and C i Tangent line l 5 ’，l 5 And l 5 ' are all indicated by two-dot chain lines, in which case 5 And l 5 The angle of' is theta 3 (ii) a Passing through point B to form circle C i Lower tangent line l 4 Then with O i To the internal load vertex B (x) load1 ,z load1 ) Is a radiusMaking an arc D 1 Arc D of a circle 1 Making a straight line l with the intersection point of the outer contour line of the machine body 4 ' and circle C i Tangent,. L 4 And l 4 ' are all indicated by one-dot chain lines, note l 4 And l 4 The angle of' is theta 2 (ii) a Then adding O i To the internal loading vertex A (x) load0 ,z load0 ) Making a circle D for the radius 2 Circle D 2 Making a straight line l with the intersection point of the outer contour line of the machine body 1 ' and circle C i Tangent,. L 1 And l 1 ' are all indicated by thin solid lines, note l 1 And l 1 The angle of' is theta 1 . Comparison of θ 1 ，θ 2 ，θ 3 The smaller is the maximum wing rotation angle, which is recorded as theta, namely theta = min (theta) 1 ,θ 2 ,θ 3 )。
Referring to FIG. 13, using C i The center of the circle is the rotation center, and the outer contour line of the machine body is connected with the line I 1 、l 2 And l 3 And rotating theta anticlockwise to obtain the shape of the wing when the wing is completely unfolded. The wing can be divided into two areas, wherein the black area consists of the part of the wing outside the machine body and a rotating shaft, and the shape (comprising an outer contour, the rotating shaft and a bottom contour) of the black area can not be changed if the wing is to be maximized; the shaded part is the rest of the wing, determines the inner contour of the wing, the part is always positioned in the body, the curve of the part can be freely selected by a designer according to the requirements of stress and the like, namely the inner contour of the wing is a self-selected area. Is a reaction with O i Correspondingly, the maximum rotation angle theta of the wing is recorded as theta i And the developed area thereof is measured as S i 。
②O i At L 2 The following steps:
referring to FIG. 14, L 2 Upper point O i The distance from the point B of the internal load is recorded as a, and a tangent circle C with the point as the center of the circle and the outer contour line z = f (x) of the machine body is formed i . Then adding O i As a circle center, taking the radius from the center to the end point N as a circular arc D, and taking a straight line z = z above the internal load at the moment load1 The intersection point with the arc D is marked as N'. Firstly, O is added i And an internally loaded vertex B (x) load1 ,z load1 ) Are connected with each other to form 1 Continue to connect the vertex BN' to l 2 。
Referring to FIG. 15, the shape of the airfoil is initially determined by a line l 1 Straight line l 2 Circular arc
Body outer contour line z = f (x), and circle C i And the enclosed black area. Namely, the wing rotating shaft, the wing outer contour and the wing bottom contour are determined.
Referring to fig. 16, in order to determine the maximum rotation angle of the adjustable wing, all the points, i.e., B and N', which may be in first contact with the outer contour line of the body are considered. First pass through N' and C i Tangent line l 3 Indicated by a dot-dash line, and then passes through the end points N and C i Tangent line l 3 ', in this case l 3 And l 3 The angle of' is theta 2 . This consideration is also needed considering that the internal loading is closer to the outer contour of the body. Therefore with O i To the internal load vertex B (x) load1 ,z load1 ) Making a circle D for the radius 1 Circle D 1 Making a straight line l with the intersection point of the outer contour line of the machine body 1 ' and circle C i Tangent, indicated by a thin solid line in the figure, note l 1 And l 1 The angle of' is theta 1 . Comparison of θ 1 And theta 2 The smaller is the maximum rotation angle of the deformation wing, which is recorded as theta, namely theta = min (theta = min) 1 ,θ 2 )。
Referring to FIG. 17, with C i The center of the circle is the rotation center, and the outer contour line of the machine body is connected with the line l 1 And l 2 And rotating the angle theta anticlockwise to obtain the shape of the deformation wing when the deformation wing is completely unfolded. The wing can be divided into two areas, wherein the black area consists of the part of the wing outside the machine body and a rotating shaft, and the shape (comprising an outer contour, the rotating shaft and a bottom contour) of the black area can not be changed if the wing is to be maximized; the shaded part is the rest part of the wing, which determines the inner contour of the wing, the part is always positioned in the body, the curve can be freely selected by a designer according to the requirements of stress and the like, namely the self-selecting area of the inner contour of the wing. Is a reaction with O i Correspondingly, the wingThe maximum rotation angle θ is recorded as θ i And the developed area thereof is measured as S i 。
③O i At L 1 The method comprises the following steps:
referring to FIG. 18, L 1 Last point O i The distance is recorded until the internal load is a, and a tangent circle C with the external contour line z = f (x) of the machine body is formed by taking the point as the center of the circle i . Then adding O i As the center of the circle, and the radius to the end point N is used as a circle D. At the moment, the upper straight line z = z of the internal load load1 Making an intersection point with the circle D, denoted as N ', and making a line passing through N' and with the circle C i Tangent line l 1 。
Referring to FIG. 19, the shape of the airfoil is initially determined as a line I 1 Circular arc
Body outer contour line z = f (x), and circle C k And the enclosed black area. Namely, the wing rotating shaft, the wing outer contour and the wing bottom contour are determined.
Referring to fig. 20, the maximum turning angle of the adjustable wing is determined, and only N' is in first contact with the outer contour line of the body. Making a pass through end points N and C i Tangent line l 1 ', when l 1 And l 1 The included angle theta of' is the maximum rotation angle of the deformation wing.
Referring to FIG. 21, using C k The center of the circle is the rotation center, and the outer contour line of the machine body is connected with the line l 1 And rotating the angle theta anticlockwise to obtain the shape of the deformation wing when the deformation wing is completely unfolded. The wing can be divided into two areas, wherein the black area consists of the part of the wing outside the machine body and a rotating shaft, and the shape (comprising an outer contour, the rotating shaft and a bottom contour) of the black area can not be changed if the wing is to be maximized; the shaded part is the rest of the wing, determines the inner contour of the wing, the part is always positioned in the body, the curve of the part can be freely selected by a designer according to the requirements of stress and the like, namely the inner contour of the wing is a self-selected area. Is a group of general formula with O i Correspondingly, the maximum rotation angle theta of the wing is recorded as theta i And the developed area thereof is measured as S i 。
And 3, selecting the one which can maximize the expansion area value from the design results of the different points as a final design result, calculating the stress of the wing according to the extreme flight condition of the aircraft, checking the stress intensity to determine the minimum thickness h which the wing should take, and finishing the design of the adjustable wing. The method comprises the following two steps:
1) According to the maximum expansion area value S k ＝max(S 0 ～S n ) Calculating the pressure difference (P) between the upper surface and the lower surface of the wing according to the extreme flight condition of the aircraft 3 -P 2 )。
Referring to FIG. 22, the Mach number M in the incoming flow I region is shown for the stress condition of the wing 1 Pressure P 1 Angle of flight attack α.
In the post-shock region III, P 3 According to the relationship θ - β -M:
the oblique shock angle beta can be obtained. According to the oblique shock wave relation:
availability of P 3 /P 1 Then the ratio can be used to obtain the pressure P in the III area 3 The value of (c).
In the post-dilation region II, the stream Mach number M can be obtained by using a Prandtl-Meyer function 1 Obtaining the Mach number M after the expansion wave 2 ：
v(M 2 )＝α+v(M 1 ) (5)
M 2 ＝v -1 [v(M 2 )] (6)
The total pressure of the gas before and after passing through the expansion wave is equal, i.e. P 01 ＝P 02 . From the basic equation of isentropic flow:
the relation between the total pressure and the static pressure after the expansion wave front can be respectively obtained, namely P 01 /P 1 And P 02 /P 2 . Then the following formula
Then P can be obtained 2 . Finally, the normal uniform pressure (P) of the whole wing is obtained 3 -P 2 )。
2) According to the normal uniform pressure (P) applied to the wing in the step 1) 3 -P 2 ) And the wing is made of a material, and the stress of the wing is checked, so that the minimum thickness h of the wing is determined.
Referring to fig. 23, since the outer contour line of the body is an irregular curve, in order to find the maximum stress position, the wing can be regarded as an infinite number of line segments perpendicular to the outer contour line of the body. After the wing is unfolded, a point Q of the outer contour of the wing, which is farthest from the body, is found, then a perpendicular line passing through Q and taking the outer contour line z = f (x) of the body is made, and the perpendicular line is intersected with Q'. The bending moment at Q' is the maximum.
Referring to FIG. 24, a small width δ b is assumed at QQ' at the maximum force. The stress can be converted into an cantilever beam type structure with the length of QQ', the width of delta b and the height of h as shown in FIG. 25. On the basis of FIG. 24, a cross-sectional view E-E is made, and the three-dimensional mechanical problem is simplified into a two-dimensional cantilever beam form; and the length of QQ' is set as l, and the height h is the minimum thickness required by the wing. The bending moment on the section at the position Q' is the maximum.
The bending moment of the section in the mechanics of materials is used for checking, and the bending section coefficient of the rectangular section with the length delta b and the width h is as follows:
W＝h 2 δb/6(10)
the maximum bending stress at Q' should then satisfy:
where [ sigma ] is the allowable bending stress of the wing material. The wing thickness h obtained from the above should satisfy:
example (b):
the present embodiment will be designed according to the above process by taking the aircraft shown in fig. 1 as an example.
Referring to fig. 26, the aircraft is half-sectioned along xOz with the top of the airframe as the origin, and its internal loading dimension is as shown, and the starting position of the air vane is the ending point N.
Referring to FIG. 27, the radius r of the rotation axis is 100mm, and the equidistant line L is drawn, and the point-drawing range is determined to be O 0 ～O n Taking multiple points O therein i As the center of the rotation shaft. The elongation lines of AB do not lie entirely on circle C 0 Below, so L can be divided into three sections L 1 ，L 2 ，L 3 And a three-stage design method is adopted.
FIG. 28 shows the design result when the center of the rotation shaft is located at O 0 The area of the stent reaches the maximum value S 0 ＝0.923m 2 Maximum deployment angle θ 0 O is used for this reason, =7.41 ° 0 The design result obtained for the center of the rotating shaft. The black area in the figure is a fixed area with an area of 0.955m 2 The shaded area is an inner contour optional area with the area of 0.377m 2 The total area of the adjustable wing is 1.340m 2 . The area between the outer contour line of the machine body and the inner load is 1.893m 2 The space occupancy rate is 50.4-70.8%A (c) is added; and the areas are all unilateral.
With the single wing spreading area S 0 And carrying out stress calculation and stress check. Assuming that the aircraft extreme operating conditions are 10Ma, an angle of attack of 15 °, a height of 30km, and a static pressure of 1197Pa. And (3) obtaining the upper and lower pressure difference of the adjustable wing when the adjustable wing is completely unfolded according to the calculation result: (P) 3 -P 2 ) =9155Pa. The farthest distance between the outer contour of the wing and the airframe when the wing is completely unfolded is measured to be l =0.486m, and the allowable bending stress of the material used by the wing is set to be [ sigma ]]And (5) =200MPa, and the thickness h of the adjustable wing is more than or equal to 5.7mm.
The final adjustable wing design result is shown in figure 1,
finally, it should be noted that: although the present invention has been described in detail with reference to the foregoing embodiments, it will be apparent to those skilled in the art that changes may be made in the embodiments and/or equivalents thereof without departing from the spirit and scope of the invention. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (8)
1. The design method of the single-rotating-shaft lift margin aircraft adjustable wing based on the geometric strong constraint is characterized by comprising the following steps of:
s1, determining a position range of a rotary shaft center according to input data, and taking a series of points as the rotary shaft center;
s2, dividing the position range into two to three areas, and obtaining corresponding design results by adopting different design processes for points in different areas;
and S3, selecting the one which can maximize the expansion area value from the design results of the different points as a final design result, calculating the stress of the wing according to the extreme flight condition of the aircraft, checking the stress intensity to determine the minimum thickness which the wing should take, and finishing the design of the adjustable wing.
2. The design method of the adjustable wing of the single-rotating-shaft lift margin aircraft based on the geometric constraint of claim 1 is characterized in that: the method for determining the position range and the point set in the S1 comprises the steps of firstly performing half-section on the aircraft along the plane where the wings are located, taking an input radius r as an inner equal-distance line L of an outer contour line of the aircraft body, enabling the center of a rotating shaft to be located on a curve L, and meeting the requirement that the gap between the outer contour line of the aircraft body and an inner load is direct for the input radius r of the rotating shaft to accommodate the rotating shaft.
3. The design method of the adjustable wing of the single-rotating-shaft lift margin aircraft based on the geometric constraint of claim 2 is characterized in that: taking n +1 points with certain intervals on the curve L as the center of the rotating shaft, and marking as { O 0 ～O n Consider, where the starting position O 0 From the central axis r, i.e. z = r, end position O n Taking distance r from the bottom surface of the circular truncated cone, namely x = x load1 -r。
4. The design method of the single-rotating-shaft lift margin aircraft adjustable wing based on the strong geometric constraint is characterized by comprising the following steps: the step of the process of dividing the position range in the step S2 and designing the position range in a sectional manner includes that the curve L is divided into two or three sections according to the geometric shape of the internal load.
5. The design method of the single-rotating-shaft lift margin aircraft adjustable wing based on the geometric constraint of claim 4 is characterized in that: designing the deformed wing according to a two-stage or three-stage method so as to determine each point O i The corresponding wing rotating shaft, the outer contour and the bottom contour of the wing, and the inner contour of the wing are initially determined.
6. The design method of the adjustable wing of the single-rotating-shaft lift margin aircraft based on the geometric constraint of claim 5 is characterized in that: for the preliminarily obtained airfoil shape to wind O i Rotating anticlockwise to obtain the result when the wing is completely unfolded and recording the completely unfolded area S i And the inner profile of the wing should be always positioned inside the machine body during the rotation process.
7. The design method of the adjustable wing of the single-rotating-shaft lift margin aircraft based on the geometric constraint of claim 1 is characterized in that: the stress and stress checking step of the wings in the step S3 comprises the step of selecting the maximum expansion area S k ＝{S 0 ～S n Calculating the pressure difference (P) of the upper surface and the lower surface of the wing according to the area and the limit flight condition of the aircraft 3 -P 2 )。
8. The design method of the adjustable wing of the single-rotating-shaft lift margin aircraft based on the geometric constraint of claim 7 is characterized in that: according to the normal uniform pressure (P) applied to the wing 3 -P 2 ) And the wing is made of a material to be used, and the stress of the wing is checked, so that the minimum thickness h of the wing is determined to meet the following conditions:
wherein l is the maximum length of the outer contour of the wing from the outer contour line of the body, and [ sigma ] is the allowable stress of the wing material.
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| US20050001103A1 (en) * | 2003-06-17 | 2005-01-06 | Vassberg John Charles | Variable trailing edge geometry and spanload control |
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| CN114313217A (en) * | 2022-01-13 | 2022-04-12 | 南京航空航天大学 | Wing capable of folding and unfolding along unfolding direction variant |
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| US20050001103A1 (en) * | 2003-06-17 | 2005-01-06 | Vassberg John Charles | Variable trailing edge geometry and spanload control |
| CN101941523A (en) * | 2010-08-12 | 2011-01-12 | 吴新保 | Adjustable aerofoil and double-body aircraft aerofoil layout scheme thereof |
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