CN107444669B - 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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Abstract
The invention discloses trans- hypersonic aircraft aerodynamic arrangement design methods under a kind of, include the following steps: specifying constraint: length L, width W, and internal diameter φ, head radius R are loaded in bottom sectionh, head spherical surface corner cut θ, nose of wing radius Rw；Step 1: determining the upper and lower surface contour line of aircraft；Step 2: determining the left and right width profile line of aircraft；Step 3: anti-cross section curve under determining；Step 4: generating the oval cross section before B point；Step 5: generating B point to the compound section between C point, aircraft shape is obtained.Quickly generating for shape under different inverted diherdrals and size constraint may be implemented in the method for the present invention, and the shape that this method generates can be with risk management, under trans- leeward both ensure that lifting area was sufficiently large, inhibit the upward spilling of windward side high pressure draught again simultaneously, reduce loss of lift, it is able to ascend pneumatic efficiency, can be designed for novel hypersonic aircraft and a kind of optional layout method and scheme newly is provided.
Description
Technical field
The present invention relates to aerodynamic arrangement's design fields of aircraft, and in particular to trans- hypersonic promotion under a kind of
Resistance is than aerodynamic configuration of aircraft design method.
Background technique
The aerodynamic arrangement of the hypersonic aircraft currently generallyd use is the flat layout for being seated in 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 will lead to it is flat
Flat fuselage lateral stability is poor, while being unfavorable for the installation of rudder face, and since inhibiting effect of the plate to lateral flow is weaker,
High pressure draught is caused to be routed up upwards, larger so as to cause pneumatic efficiency loss, lift resistance ratio is smaller.In addition, although hypersonic fly
Hang Qi aerodynamic arrangement has 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 bending angle.Therefore in order to further enhance hypersonic aircraft aeroperformance, gram
Take its lateral stability there are the problem of, need to explore new aerodynamic arrangement's form.
Summary of the invention
In order to overcome the disadvantages mentioned above of the prior art, the invention proposes the trans- pneumatic cloth of hypersonic aircraft under a kind of
Office's design method is seated in top lifting surface in the distribution form of lower part by changing conventional aircraft, filling is placed in
Lower part, trans- planar design under leeward surface curve uses, similar to the bionical shape of bat.The benefit of this design is to meet
Trans- leeward not only ensure that sufficiently large lifting area under under the premise of load requirement, but also inhibit windward side high pressure gas
The upward spilling of stream, reduces loss of lift, improves pneumatic efficiency, can also be the design of rudder face with leeward trans- at present
It is convenient to provide.
The technical solution adopted by the present invention to solve the technical problems is: the trans- pneumatic cloth of hypersonic aircraft under a kind of
Office's design method includes the following steps: the following constraint condition of given aircraft design: length L, width W, bottom section filling
Internal diameter φ, head radius Rh, head spherical surface corner cut θ, nose of wing radius Rw；
Step 1: determining the upper and lower surface contour line of aircraft: upper surface contour line is that circular arc AB adds conic section BC, circle
The central coordinate of circle of arc AB is (xo=Rh,zo=0), A point is coordinate origin, and B point coordinate is (xB=Rh-Rh* sin θ, zB=Rh*cos
θ), C point coordinate is (xC=L, zC=φ), lower surface contour line and upper surface contour line are symmetrical about x-axis；
Step 2: determining 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 point is (xB1=Rh-Rh* sin θ, yB1=Rh* cos θ), the seat of C1
It is designated as (xC1=L, yC1=W/2), lower surface contour line and upper surface contour line are symmetrical about x-axis；
Step 3: anti-cross section curve under determining: under the radius of the anti-cross section curve circle of dotted line that includes be equal to bottom section dress
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 coordinate be (xD=L, yD=φ sin θ 1, zD=φ cos θ 1), straight line DE respectively with circular arc C D
Tangent with circular arc EC1F, circular arc EC1F radius is Rw, central coordinate of circle be (xO1=L, yO1=W/2-Rw, zO1=zE-Rw*cos(θ
1)), E point coordinate 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 θ 1)), F point coordinate is (xF=L, yF=yo1-Rw* sin (θ 2), zF=zO1-Rw* cos (θ 2)), G point coordinate 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 point, H point
Coordinate is (xH=L, yH=W/6, zH=-0.4 φ), I point coordinate is (xI=L, yI=0, zI=-φ)；Left and right half portion bottom is cut
The curve in face is symmetrical about Z axis；
Step 4: generating the oval cross section before B point；
Step 5: generating B point to the compound section between C point, aircraft shape is obtained.
Compared with prior art, the positive effect of the present invention is: different inverted diherdrals and size may be implemented in the method for the present invention
Shape quickly generates under constraint condition, and the shape that generates of this method can with risk management, under trans- leeward both protected
It has demonstrate,proved that lifting area is sufficiently large, while having 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 novel hypersonic aircraft and a kind of optional layout method and scheme newly is provided.
Detailed description of the invention
Examples of the present invention will be described by way of reference to the accompanying drawings, in which:
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 the different sections x；
Fig. 6 is the outline drawing ultimately generated.
Specific embodiment
Trans- hypersonic aircraft aerodynamic arrangement design method, includes the following steps: under a kind of
Step 1: internal diameter φ, head radius R are loaded in the length L of given aircraft design, width W, bottom sectionh, head
Spherical surface corner cut θ, nose of wing radius RwConstraint condition.The upper and lower surface contour line of aircraft, upper surface are determined according to constraint condition
Contour line is that head radius is RhCircular arc add conic section, arc section corresponding angle is 90 ° of-θ, with conic section with corner cut θ
Tangent, specific curve form is shown in that Fig. 1, A point are coordinate origin, and B point is arc section end point, and BC is conic section, circular arc and two
The point of contact of secondary curve is B, according to head radius RhWith 90 ° of-θ of arc angle, B point coordinate is calculated are as follows:
It is available with O (x thus according to A, B point coordinateo=Rh,zoIt=0) is the center of circle, with RhFor the circular arc AB of radius.
Simultaneously according to length L and bottom section filling internal diameter φ constraint, available C point coordinate is (xC=L, zC=φ),
Known B point slope is tan θ, and aircraft tail portion is more gentle, and giving C point slope is 0, by the coordinate and slope of B, C two o'clock, is used
Two o'clock line can be obtained contour line by the conic section that the raised factor is 0.01, 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, internal diameter φ, head radius R are loaded in bottom sectionhIt determines
Left and right width profile line, design method are shown in Fig. 2, and wherein A point is coordinate origin, and B1 point is equally by head radius RhWith head spherical surface
Coordinate is calculated in corner cut θ:
It is available with O (x thus according to A, B1 point coordinateo=Rh,yoIt=0) is the center of circle, with RhFor the circular arc AB1 of radius.
It is constrained simultaneously according to length L and width W, available point C1 coordinate is (xC1=L, yC1=W/2), between B1, C1
It is connected with straight line, available straight line B1C1, thus completes width profile line constrained designs, left and right width profile line is about x-axis
It is full symmetric.
Step 3: anti-cross section curve under design, as shown in Figure 3 and Figure 4:
Circle of dotted line CO'I is to load the filling that internal diameter φ is determined by bottom section to constrain, and O' coordinate is on the section bottom x=L
Origin (xo'=L, yo'=0, zo'=0), lower anti-cross section curve must include the constraint.
Under anti-cross section curve be made of circular arc C D, straight line DE, circular arc EC1F, conic section FG and HI, each curve joint
Slope rate continuity.
The corresponding angle theta 1 of circular arc C D is inverted diherdral degree, according to the radius φ of CO'I, the expression of available arc section CD
Formula are as follows:
It is hereby achieved that D point coordinate (xD=L, yD=φ sin θ 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 maximum width contour line determines, coordinate yC1=W/2, therefore available yo1=W/2-Rw,
yE=W/2-Rw+Rw* sin (θ 1), the expression formula of straight line DE are as follows:
By yEBring the straight line expression formula into, available E point coordinate:
(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 θ 1)) it is hereby achieved that O1 point coordinate are as follows:
It is possible thereby to completely obtain the expression formula of circular arc EC1F:
Wherein θ 2 is the angle of lower aerofoil FG and horizontal direction, as shown in figure 4, obtaining F point coordinate in turn:
(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 the expression formula of straight line FG are as follows:
Take yG=W/4 obtains G point coordinate:
(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 point, loads internal diameter φ constraint according to bottom section, I point is sat
It is designated as:
(xI=L, yI=0, zI=-φ)
Take H point coordinate are as follows: (xH=L, yH=W/6, zH=-0.4 φ), it is K that H point, which gives slope,H, then with protrusion between IH
The conic section of the factor 0.5 connects, and HG is also the connection of 0.5 conic section with the raised factor, so far completes right side bottom and cuts
The curve shape in face designs, by it about the symmetrical available entire bottom section curve shape of Z axis, as shown in Figure 3.
Step 4: the head section before B point is ellipse, in arbitrary section Xi(Xi≤xB) at, it, can be with according to step 1
Obtain its corresponding upper contour line constraint Zi, according to step 2, its available corresponding width profile line constrains Yi, thus may be used
To obtain the expression formula of oval cross section:
Step 5: to the section of C point for gradually from elliptical light degree of slipping over to bottom section, section form is after B point
Oval and bottom section linear combination, for the arbitrary section X between B point and C pointi(xB≤Xi≤xC), according to step 1,
Its available corresponding upper contour line constrains Zi, according to step 2, its available corresponding width profile line constrains Yi, root
It is constrained according to the contour line, the expression formula of available elliptical profile:
The y of each point on the bottom section the x=L curve that step 4 is obtainedLAnd zLCoordinate is contracted as follows respectively
It puts, it is available to meet XiThe curvilinear coordinate form of width profile line and the constraint of upper lower contour at section:
y2=yL*Yi/ W, z2=zL*Zi/φ (9)
Each section X between C point is arrived after B pointi(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 for formula (8) acquiring oval cross section is substituted into respectively1And generation
Enter the coordinate z of bottom section curve after formula (9) scale2, new curvilinear coordinate is the linear combination of the two:
So far Curve Design at all sections x is completed, trans- shape under this is generated.
The concrete application of the method for the present invention 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 spherical surface corner cut θ=3.5 °, nose of wing radius Rw=15mm constraint condition.According to constraint
Condition determines that the upper and lower surface contour line of aircraft, upper surface contour line are that head radius is RhThe circular arc of=20mm adds secondary song
Line, arc section corresponding angle be 86.5 °, it is tangent with 3.5 ° of corner cut with conic section, available A point be coordinate origin (0,0,
0), B point is arc section end point, and coordinate is (18.78,0,19.96), and B point tangent slope is that the coordinate of 0.0611, C point is
(5000,0,275), slope 0.Thus AB sections are R with radiushThe circular sliding slopes of=20mm, BC sections of known coordinates and slope can
It is the connection of 0.01 conic section by the raised factor.Curve is obtained as shown in Figure 1, each available xiLocate corresponding height about
Beam Zi。
B1 point coordinate (x is similarly calculatedB1=18.78, yB1=19.96), C1 point coordinate (xC1=5000, yC1=
900), AB1 sections are R with radiushThe circular sliding slopes of=20mm, B1C1 sections are connected with straight line, as shown in Fig. 2, available each
A xiLocate corresponding wide constraint Yi。
D point coordinate is calculated according to step 3 in known φ=275mm, given inverted diherdral θ 1=25 °, θ 2=15 °
(5000,116.22,249.23), E point coordinate be (5000,891.34, -112.21), O1 point coordinate be (5000,885, -
125.80), F point coordinate is (5000,881.12, -145.52), and C1 point coordinate is (5000,900, -125.80), and G point coordinate is
(5000,450, -30), H point coordinate are (5000,300, -110), give conic section slope KH=1, I point coordinate be (5000,
0, -275) conic section slope 0, is given, it is known that above-mentioned coordinate and slope use circular arc, straight line, circular arc, straight line, secondary respectively
Curve, which is attached, can be obtained cross section curve shown in Fig. 3.
After completing bottom section design, according to step 4, the oval cross section before B point is generated, B point is generated according to step 5
Compound section between C point, Fig. 5 give the cross section curve shape at different location, and Fig. 6 gives finally by these sections
The aircraft shape of composition.
Claims (5)
1. trans- hypersonic aircraft aerodynamic arrangement design method under a kind of, characterized by the following steps: given to fly
The following constraint condition of row device design: internal diameter φ, head radius R are loaded in length L, width W, bottom sectionh, head spherical surface corner cut
θ, nose of wing radius Rw；
Step 1: determining 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 point is coordinate origin, and B point coordinate is (xB=Rh-Rh* sin θ, zB=Rh*cosθ)、C
Point coordinate is (xC=L, zC=φ), lower surface contour line and upper surface contour line are symmetrical about x-axis；
Step 2: determining 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 point is (xB1=Rh-Rh* sin θ, yB1=Rh* cos θ), the coordinate of C1 is
(xC1=L, yC1=W/2), right width profile line and left width profile line are symmetrical about y-axis；
Step 3: anti-cross section curve under determining: under the radius of the anti-cross section curve circle of dotted line that includes be equal in the filling of bottom section
Diameter 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, the corresponding folder of circular arc C D
Angle θ 1 is inverted diherdral degree, and D point coordinate is (xD=L, yD=φ sin θ 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 radius is Rw, central coordinate of circle be (xO1=L, yO1=W/2-Rw, zO1=zE-Rw* cos (θ 1)), E
Point coordinate is (xE=L, yE=W/2-Rw+Rw* sin (θ 1), zE=φ * cos (θ 1)-tan (θ 1) * (W/2-Rw+Rw*sin(θ1)-
φ * sin θ 1)), F point coordinate is (xF=L, yF=yo1-Rw* sin (θ 2), zF=zO1-Rw* cos (θ 2)), G point coordinate 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 point, and H point coordinate is
(xH=L, yH=W/6, zH=-0.4 φ), I point coordinate is (xI=L, yI=0, zI=-φ)；The song of left and right half portion bottom section
Line is symmetrical about Z axis；
Step 4: generating the oval cross section before B point；
Step 5: generating B point to the compound section between C point, aircraft shape is obtained.
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 B, C two o'clock line 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 described in step 3 are all made of the conic section line that the raised factor is 0.5 and obtain.
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 the point of B described in step 4i(Xi≤xB) at oval cross section expression formula are as follows:
Wherein: YiFor the constraint of width profile line, ZiFor the constraint of upper lower contour.
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:
Method of the generation B point to the compound section between C point described in step 5 are as follows:
(1) the arbitrary section X between B point and C point is determinedi(xB≤Xi≤xC) at elliptical profile expression formula are as follows:
Wherein: YiFor the constraint of width profile line, ZiFor the constraint of upper lower contour；
(2) y of each point on the bottom section the x=L 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) and the scaling formula of (2) are substituted into respectively, are asked
Obtain the coordinate z of oval cross section1With the coordinate z of bottom section curve after scaling2, obtain meeting XiWidth profile line and upper at section
The curvilinear coordinate of lower contour constraint:
Wherein, α=(2.0* ((xi-xB)/(xC-xB))3-3.0*((xi-xB)/(xC-xB))2)+1.0。
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| CN110127076A (en) * | 2019-05-22 | 2019-08-16 | 中国空气动力研究与发展中心 | A kind of flat roof type Horizontal Take-off and Landing two-stage enters orbit aerocraft level-one aerodynamic arrangement design method |
| CN112016164B (en) * | 2020-09-09 | 2022-07-01 | 中国空气动力研究与发展中心计算空气动力研究所 | Aerospace model flight test layout, and axisymmetric nose cone region configuration and design method thereof |
| CN113139243B (en) * | 2021-06-22 | 2021-08-20 | 中国空气动力研究与发展中心计算空气动力研究所 | Standard model layout design method suitable for shock wave/boundary layer interference and heat protection research |
| CN113239473B (en) * | 2021-07-12 | 2021-09-21 | 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 |
| CN113850032B (en) * | 2021-12-02 | 2022-02-08 | 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 |
| 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 |
| 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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| US20080105783A1 (en) * | 2006-11-03 | 2008-05-08 | Cessna Aircraft Company | Fuselage design for sonic boom suppression of supersonic aircraft |
| CN105173116B (en) * | 2015-09-25 | 2017-03-29 | 北京航空航天大学 | The close curved surface Waverider method for designing of hypersonic aircraft |
| US10953979B2 (en) * | 2015-11-11 | 2021-03-23 | The Arizona Board Of Regents On Behalf Of The University Of Arizona | Control of hypersonic boundary layer transition |
| CN106156403B (en) * | 2016-06-21 | 2019-07-09 | 南京航空航天大学 | Hypersonic aircraft spar structure analysis method for reliability |
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