CN107444669A - Trans hypersonic aircraft aerodynamic arrangement design method under a kind of - Google Patents
Trans hypersonic aircraft aerodynamic arrangement design method under a kind of Download PDFInfo
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- CN107444669A CN107444669A CN201710638514.3A CN201710638514A CN107444669A CN 107444669 A CN107444669 A CN 107444669A CN 201710638514 A CN201710638514 A CN 201710638514A CN 107444669 A CN107444669 A CN 107444669A
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Abstract
The invention discloses trans hypersonic aircraft aerodynamic arrangement design method under a kind of, comprise the following steps：Specifying constraint：Length L, width W, bottom section filling internal diameter φ, head radius Rh, head sphere corner cut θ, nose of wing radius Rw；Step 1: determine the upper and lower surface contour line of aircraft；Step 2: determine the left and right width profile line of aircraft；Step 3: anti-cross section curve under determining；Step 4: the oval cross section before generation B points；Step 5: generation B points obtain aircraft profile to the compound section between C points.The inventive method can realize quickly generating for profile under different inverted diherdrals and size constraint, and the profile of this method generation can be with risk management, under trans lee face both ensure that lifting area was sufficiently large, inhibit the upward spilling of windward side high pressure draught again simultaneously, reduce loss of lift, pneumatic efficiency can be lifted, can be designed for new hypersonic aircraft and a kind of new optional layout method and scheme are provided.
Description
Technical field
The present invention relates to aerodynamic arrangement's design field of aircraft, and in particular to trans hypersonic promotion under a kind of
Resistance is than aerodynamic configuration of aircraft design method.
Background technology
The aerodynamic arrangement of the current hypersonic aircraft generally used to be seated in the flat layout on top, i.e., bottom compared with
For gentle blended wing-body or assembly configuration.But this method has disadvantages that, as windward side flat sheet configuration can cause it is flat
Flat fuselage lateral stability is poor, while is unfavorable for the installation of rudder face, and because flat board is weaker to the inhibitory action of lateral flow,
High pressure draught is caused to be routed up upwards, so as to cause pneumatic efficiency loss larger, lift-drag ratio is smaller.In addition, although hypersonic fly
Hang Qi aerodynamic arrangements have achieved a series of achievements, but its be laid out cross sectional shape still concentrate on it is traditional be seated in top, under
Portion is plane or the plane with certain angle of bend.Therefore in order to further lift hypersonic aircraft aeroperformance, gram
The problem of its lateral stability is present is taken, it is necessary to explore new aerodynamic arrangement's form.
The content of the invention
In order to overcome the disadvantages mentioned above of prior art, the present invention propose it is a kind of under the trans pneumatic cloth of hypersonic aircraft
Office's design method, top is seated in and distribution form of the lifting surface in bottom by changing conventional aircraft, filling is positioned over
Bottom, trans planar design under leeward surface curve uses, similar to the bionical shape of bat.The benefit of this design is to meet
Trans lee face both ensure that sufficiently large lifting area under on the premise of load requirement, while inhibit windward side high pressure gas again
The upward spilling of stream, reduces loss of lift, improves pneumatic efficiency, can also be the design of rudder face with lee face trans at present
It is convenient to provide.
The technical solution adopted for the present invention to solve the technical problems is：The trans pneumatic cloth of hypersonic aircraft under a kind of
Office's design method, comprises the following steps：The following constraints of given aircraft design：Length L, width W, bottom section filling
Internal diameter φ, head radius Rh, head sphere corner cut θ, nose of wing radius Rw；
Step 1: determine the upper and lower surface contour line of aircraft：Upper surface contour line is that circular arc AB adds conic section BC, circle
Arc AB central coordinate of circle is (xo=Rh,zo=0), A points are the origin of coordinates, and B point coordinates is (xB=Rh-Rh* sin θ, zB=Rh*cos
θ), C point coordinates is (xC=L, zC=φ), lower surface contour line and upper surface contour line are symmetrical on x-axis；
Step 2: determine the left and right width profile line of aircraft：Left width profile line is that circular arc AB1 adds straightway B1C1,
The central coordinate of circle of circular arc is (xo=Rh,yo=0), the coordinate of B1 points is (xB1=Rh-Rh* sin θ, yB1=Rh* cos θ), C1 seat
It is designated as (xC1=L, yC1=W/2), lower surface contour line and upper surface contour line are symmetrical on x-axis；
Step 3: anti-cross section curve under determining：Under the radius of broken circle that includes of anti-cross section curve filled equal to bottom section
Fill out internal diameter φ, central coordinate of circle is (xo'=L, yo'=0, zo'=0)；Under anti-cross section curve right side bottom section curve by
Circular arc C D, straight line DE, circular arc EC1F, conic section FG, GH and HI composition, each curve joint slope rate continuity, circular arc C D are corresponding
Angle theta 1 be inverted diherdral degree, D point coordinates is (xD=L, yD=φ sin θs 1, zD=φ cos θ 1), straight line DE respectively with circular arc C D
Tangent with circular arc EC1F, circular arc EC1F radiuses are Rw, central coordinate of circle be (xO1=L, yO1=W/2-Rw, zO1=zE-Rw*cos(θ
1)), E point coordinates is (xE=L, yE=W/2-Rw+Rw* sin (θ 1), zE=φ * cos (θ 1)-tan (θ 1) * (W/2-Rw+Rw*sin
(θ 1)-φ * sin θs 1)), F point coordinates is (xF=L, yF=yo1-Rw* sin (θ 2), zF=zO1-Rw* cos (θ 2)), G point coordinates is
(xG=L, yG=W/4, zG=zF-tan(θ2)*(yG-yF)), GH and HI are that two sections of splicing conic sections are tangent at H points, H points
Coordinate is (xH=L, yH=W/6, zH=-0.4 φ), I point coordinates is (xI=L, yI=0, zI=-φ)；Left and right half portion bottom is cut
The curve in face is symmetrical on Z axis；
Step 4: the oval cross section before generation B points；
Step 5: generation B points obtain aircraft profile to the compound section between C points.
Compared with prior art, the positive effect of the present invention is：The inventive method can realize different inverted diherdrals and size
Profile quickly generates under constraints, and the profile of this method generation can with risk management, under trans lee face both protected
Demonstrate,prove that lifting area is sufficiently large, while inhibited the upward spilling of windward side high pressure draught again, reduced loss of lift, Neng Gouti
Pneumatic efficiency is risen, can be designed for new hypersonic aircraft and a kind of new optional layout method and scheme are provided.
Brief description of the drawings
Examples of the present invention will be described by way of reference to the accompanying drawings, wherein：
Fig. 1 is upper and lower surface contour line design result；
Fig. 2 is left and right width profile line design result；
Fig. 3 is bottom section contour line design result；
Fig. 4 is the enlarged drawing of circular arc EC1F in Fig. 3；
Fig. 5 is Curve Design result at different x sections；
Fig. 6 is the outline drawing ultimately generated.
Embodiment
Trans hypersonic aircraft aerodynamic arrangement design method, comprises the following steps under a kind of：
Step 1：The length L of given aircraft design, width W, bottom section filling internal diameter φ, head radius Rh, head
Sphere corner cut θ, nose of wing radius RwConstraints.The upper and lower surface contour line of aircraft, upper surface are determined according to constraints
Contour line is that head radius is RhCircular arc add conic section, arc section corresponding angle is 90°- θ, with conic section with corner cut θ phases
Cut, specific curve form is shown in Fig. 1, and A points are the origin of coordinates, and B points are arc section end point, and BC is conic section, circular arc with it is secondary
The point of contact of curve is B, according to head radius RhWith arc angle 90°- θ, B point coordinates, which is calculated, is：
It can be obtained with O (x thus according to A, B point coordinateso=Rh,zo=0) it is the center of circle, with RhFor the circular arc AB of radius.
Internal diameter φ constraints are loaded according to length L and bottom section simultaneously, it is (x that can obtain C point coordinatesC=L, zC=φ),
Known B points slope is tan θ, and aircraft afterbody is more gentle, and it is 0 to give C points slope, by 2 points of coordinate and slope of B, C, is used
The conic section that the raised factor is 0.01 can obtain 2 lines contour line, and lower surface contour line and upper surface contour line close
It is full symmetric in x-axis.
Step 2：According to the length L of Flight Vehicle Design, width W, bottom section filling internal diameter φ, head radius RhIt is determined that
Left and right width profile line, design method are shown in Fig. 2, and wherein A points are the origin of coordinates, and B1 points are equally by head radius RhWith head sphere
Coordinate is calculated in corner cut θ：
It can be obtained with O (x thus according to A, B1 point coordinateso=Rh,yo=0) it is the center of circle, with RhFor the circular arc AB1 of radius.
Constrained simultaneously according to length L and width W, it is (x that can obtain point C1 coordinatesC1=L, yC1=W/2), between B1, C1
Connected with straight line, straight line B1C1 can be obtained, thus complete width profile line constrained designs, left and right width profile line is on x-axis
It is full symmetric.
Step 3：Anti- cross section curve under design, as shown in Figure 3 and Figure 4：
Broken circle CO'I is to be constrained by the bottom section filling internal diameter φ fillings determined, and O' coordinates are on the x=L sections of bottom
Origin (xo'=L, yo'=0, zo'=0), lower anti-cross section curve must include the constraint.
Under anti-cross section curve be made up of circular arc C D, straight line DE, circular arc EC1F, conic section FG and HI, each curve joint
Slope rate continuity.
Angle theta 1 corresponding to circular arc C D is inverted diherdral degree, according to CO'I radius φ, can obtain arc section CD expression
Formula is：
It is hereby achieved that D point coordinates (xD=L, yD=φ sin θs 1, zD=φ cos θ 1).
Straight line DE and circular arc C D are tangent, while tangent with circular arc EC1F, and circular arc EC1F is by nose of wing radius RwConstraint determination,
Its radius is Rw, C1 is the point that Breadth Maximum contour line determines, its coordinate is yC1=W/2, therefore y can be obtainedo1=W/2-Rw,
yE=W/2-Rw+Rw* sin (θ 1), straight line DE expression formula are：
By yEBring the straight line expression formula into, E point coordinates can be obtained：
(xE=L, yE=W/2-Rw+Rw*sin(θ1),zE=φ * cos (θ 1)-tan (θ 1) * (W/2-Rw+Rw*sin(θ1)-
φ * sin θs 1)) it is hereby achieved that the coordinate of O1 points is：
It is possible thereby to intactly obtain circular arc EC1F expression formula：
Wherein θ 2 is the angle of lower aerofoil FG and horizontal direction, as shown in figure 4, and then obtaining F point coordinates：
(xF=L, yF=yo1-Rw*sin(θ2),zF=zO1-Rw*cos(θ2))
FG is the straight line tangent with circular arc EC1F, and slope is-tan (θ 2), therefore straight line FG expression formula is：
Take yG=W/4, obtain G point coordinates：
(xG=L, yG=W/4, zG=zF-tan(θ2)*(yG-yF))
IH and HG is that two sections of splicing conic sections are tangent at H points, loads internal diameter φ constraints according to bottom section, I points are sat
It is designated as：
(xI=L, yI=0, zI=-φ)
The H point coordinates is taken to be：(xH=L, yH=W/6, zH=-0.4 φ), it is K that H points, which give slope,H, then with projection between IH
The conic section connection of the factor 0.5, HG are also connected with the raised factor for 0.5 conic section, are so far completed right side bottom and are cut
The curve shape design in face, can symmetrically obtain whole bottom section curve shape on Z axis, as shown in Figure 3 by it.
Step 4：Head section before B points is ellipse, in arbitrary section Xi(Xi≤xB) place, can be with according to step 1
Obtain upper contour line constraint Z corresponding to iti, according to step 2, its corresponding width profile line constraint Y can be obtainedi, thus may be used
To obtain the oval expression formula in section：
Step 5：To the section of C points for progressively from elliptical light degree of slipping over to bottom section, its section form is after B points
Oval and bottom section linear combination, for the arbitrary section X between B points and C pointsi(xB≤Xi≤xC), according to step 1,
Upper contour line constraint Z corresponding to it can be obtainedi, according to step 2, its corresponding width profile line constraint Y can be obtainedi, root
Constrained according to the contour line, the expression formula of elliptical profile can be obtained：
The y of each point on the x=L bottom sections curve that step 4 is obtainedLAnd zLCoordinate is contracted as follows respectively
Put, X can be metiWidth profile line and the curvilinear coordinate form of upper lower contour constraint at section：
y2=yL*Yi/ W, z2=zL*Zi/φ (9)
To each section X between C points after B pointsi(xB≤Xi≤xC) curve representation formula be above-mentioned two formula combination,
Specific combining form is for identical yi(-Yi≤yi≤Yi), the coordinate z that formula (8) tries to achieve oval cross section is substituted into respectively1And generation
Enter the coordinate z of bottom section curve after formula (9) scales2, new curvilinear coordinate is the linear combination of the two：
So far Curve Design at all x sections is completed, generates trans profile under this.
The concrete application of the inventive method illustrated below：
The length L=5000mm of given aircraft design, width W=1800mm, bottom section filling internal diameter φ=
275mm, head radius Rh=20mm, head sphere corner cut θ=3.5 °, nose of wing radius Rw=15mm constraintss.According to constraint
Condition determines the upper and lower surface contour line of aircraft, and upper surface contour line is that head radius is Rh=20mm circular arc adds secondary song
Line, arc section corresponding angle be 86.5 °, it is tangent with 3.5 ° of corner cut with conic section, can obtain A points for the origin of coordinates (0,0,
0), B points are arc section end point, and its coordinate is (18.78,0,19.96), and B points tangent slope is for the coordinate of 0.0611, C points
(5000,0,275), slope 0.Thus AB sections are R with radiush=20mm circular sliding slopes, BC sections known coordinate and slope can
Connected by the raised factor for 0.01 conic section.Curve is obtained as shown in figure 1, each x can be obtainediHeight corresponding to place is about
Beam Zi。
B1 point coordinates (x is similarly calculatedB1=18.78, yB1=19.96), C1 point coordinates (xC1=5000, yC1=
900), AB1 sections are R with radiush=20mm circular sliding slopes, B1C1 sections are connected with straight line, as shown in Fig. 2 can obtain each
Individual xiWide constraint Y corresponding to placei。
Known φ=275mm, inverted diherdral θ 1=25 °, θ 2=15 ° are given, according to step 3, D point coordinates is calculated
(5000,116.22,249.23), E point coordinates are (5000,891.34, -112.21), O1 point coordinates for (5000,885, -
125.80), F point coordinates is (5000,881.12, -145.52), and C1 point coordinates is (5000,900, -125.80), and G point coordinates is
(5000,450, -30), H point coordinates are (5000,300, -110), give conic section slope KH=1, I point coordinates for (5000,
0, -275) conic section slope 0, is given, it is known that above-mentioned point coordinates and slope are respectively with circular arc, straight line, circular arc, straight line, secondary
Curve is attached the cross section curve that can obtain shown in Fig. 3.
After completing bottom section design, according to step 4, the oval cross section before B points is generated, B points are generated according to step 5
Compound section between C points, Fig. 5 give the cross section curve shape at diverse location, and Fig. 6 is given finally by these sections
The aircraft profile of composition.
Claims (5)
1. trans hypersonic aircraft aerodynamic arrangement design method under a kind of, it is characterised in that：Comprise the following steps：It is given to fly
The following constraints of row device design：Length L, width W, bottom section filling internal diameter φ, head radius Rh, head sphere corner cut
θ, nose of wing radius Rw；
Step 1: determine the upper and lower surface contour line of aircraft：Upper surface contour line is that circular arc AB adds conic section BC, circular arc AB
Central coordinate of circle be (xo=Rh,zo=0), A points are the origin of coordinates, and B point coordinates is (xB=Rh-Rh* sin θ, zB=Rh*cosθ)、C
Point coordinates is (xC=L, zC=φ), lower surface contour line and upper surface contour line are symmetrical on x-axis；
Step 2: determine the left and right width profile line of aircraft：Left width profile line is that circular arc AB1 adds straightway B1C1, circular arc
Central coordinate of circle be (xo=Rh,yo=0), the coordinate of B1 points is (xB1=Rh-Rh* sin θ, yB1=Rh* cos θ), C1 coordinate is
(xC1=L, yC1=W/2), lower surface contour line and upper surface contour line are symmetrical on x-axis；
Step 3: anti-cross section curve under determining：Under the radius of broken circle that includes of anti-cross section curve be equal in the filling of bottom section
Footpath φ, central coordinate of circle are (xo'=L, yo'=0, zo'=0)；Under anti-cross section curve right side bottom section curve by circular arc
CD, straight line DE, circular arc EC1F, conic section FG, GH and HI composition, each curve joint slope rate continuity, folder corresponding to circular arc C D
Angle θ 1 is inverted diherdral degree, and D point coordinates is (xD=L, yD=φ sin θs 1, zD=φ cos θ 1), straight line DE respectively with circular arc C D and circle
Arc EC1F is tangent, and circular arc EC1F radiuses are Rw, central coordinate of circle be (xO1=L, yO1=W/2-Rw, zO1=zE-Rw* cos (θ 1)), E
Point coordinates is (xE=L, yE=W/2-Rw+Rw*sin(θ1), zE=φ * cos (θ 1)-tan (θ 1) * (W/2-Rw+Rw*sin(θ1)-
φ * sin θs 1)), F point coordinates is (xF=L, yF=yo1-Rw* sin (θ 2), zF=zO1-Rw* cos (θ 2)), G point coordinates is (xG=
L, yG=W/4, zG=zF-tan(θ2)*(yG-yF)), GH and HI are that two sections of splicing conic sections are tangent at H points, and H point coordinates is
(xH=L, yH=W/6, zH=-0.4 φ), I point coordinates is (xI=L, yI=0, zI=-φ)；The song of left and right half portion bottom section
Line is symmetrical on Z axis；
Step 4: the oval cross section before generation B points；
Step 5: generation B points obtain aircraft profile to the compound section between C points.
2. it is according to claim 1 it is a kind of under trans hypersonic aircraft aerodynamic arrangement design method, it is characterised in that：
Conic section BC described in step 1 uses the raised factor to obtain 2 lines of B, C for 0.01 conic section.
3. it is according to claim 1 it is a kind of under trans hypersonic aircraft aerodynamic arrangement design method, it is characterised in that：
Conic section GH and HI use the raised factor to be obtained for 0.5 conic section line described in step 3.
4. it is according to claim 1 it is a kind of under trans hypersonic aircraft aerodynamic arrangement design method, it is characterised in that：
Arbitrary section X before B points described in step 4i(Xi≤xB) the oval cross section expression formula at place is：
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5. it is according to claim 4 it is a kind of under trans hypersonic aircraft aerodynamic arrangement design method, it is characterised in that：
B points are generated described in step 5 is to the method for the compound section between C points：
(1) the arbitrary section X between B points and C points is determinedi(xB≤Xi≤xC) the elliptical profile expression formula at place is：
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Wherein：YiConstrained for width profile line, ZiConstrained for upper lower contour；
(2) y of each point on the x=L bottom sections curve for obtaining step 4LAnd zLCoordinate is scaled as follows respectively：
y2=yL*Yi/ W, z2=zL*Zi/φ；
(3) for identical yi(-Yi≤yi≤Yi), the elliptical profile expression formula of (1) step and the scaling of (2) step are substituted into respectively
Formula, try to achieve the coordinate z of oval cross section1With the coordinate z of bottom section curve after scaling2, it is met XiWidth profile at section
Line and the curvilinear coordinate of upper lower contour constraint：
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Wherein, α=(2.0* ((xi-xB)/(xC-xB))3-3.0*((xi-xB)/(xC-xB))2)+1.0。
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| CN113239473A (en) * | 2021-07-12 | 2021-08-10 | 中国空气动力研究与发展中心计算空气动力研究所 | Lifting body standard die design method for composite material performance prediction and aircraft |
| CN113850032A (en) * | 2021-12-02 | 2021-12-28 | 中国空气动力研究与发展中心计算空气动力研究所 | Load balancing method in numerical simulation calculation |
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| CN113239473A (en) * | 2021-07-12 | 2021-08-10 | 中国空气动力研究与发展中心计算空气动力研究所 | Lifting body standard die design method for composite material performance prediction and aircraft |
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| CN114117648A (en) * | 2022-01-24 | 2022-03-01 | 中国空气动力研究与发展中心计算空气动力研究所 | Combined passive control structure for simultaneously inhibiting Mack mode and transverse flow instability |
| CN114117648B (en) * | 2022-01-24 | 2022-04-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Combined passive control structure for simultaneously inhibiting Mack mode and transverse flow instability |
| CN117382898A (en) * | 2023-12-08 | 2024-01-12 | 中国空气动力研究与发展中心计算空气动力研究所 | Construction method of pneumatic layout of power energy-supplementing bouncing gliding lifting body |
| CN117382898B (en) * | 2023-12-08 | 2024-02-20 | 中国空气动力研究与发展中心计算空气动力研究所 | Construction method of pneumatic layout of power energy-supplementing bouncing gliding lifting body |
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