CN105667812A - Waverider integration design method for hypersonic aircraft forebody, air inlet and wing - Google Patents

Abstract

The invention provides a waverider integration design method for a hypersonic aircraft forebody, an air inlet and a wing. On the basis of waverider forebody/air inlet integration design, the theory of characteristic lines is applied for establishing a hypersonic aircraft forebody-air inlet integration axial symmetrical benchmark flow field and a wing symmetrical benchmark flow field, and a hypersonic aircraft forebody-air inlet integration configuration and a wing configuration are generated by applying the streamline-tracing technique in the two benchmark flow fields correspondingly. The two configurations together form a hypersonic aircraft forebody-air inlet-wing waverider integration configuration. In the design state, a whole hypersonic aircraft external flow field has waverider-like characteristics, the waverider forebody is used as a precompression face of the air inlet, and precompressed airflow is efficiently captured and provided for the air inlet, so that the waverider wing provides a high lift-drag ratio for the aircraft.

Description

Hypersonic aircraft precursor, air intake duct and wing rider integrated design method

Technical field

The present invention relates to the technical field of Air-breathing hypersonic vehicle Design of Aerodynamic Configuration, be specifically related to a kind of hypersonic aircraft precursor, air intake duct and wing rider integrated design method.

Background technology

Air-breathing hypersonic vehicle refer to flight Mach number more than 5, with airbreathing motor or its combined engine be major impetus, can at atmosphere with across the aircraft of atmosphere medium-long range flight, its application form includes hypersonic cruise missile, hypersonic has multiple aircraft such as people/unmanned aerial vehicle and sky and space plane etc.

Big quantity research since the sixties in 20th century absolutely proves, the integrated design of propulsion system and body is the key realizing hypersonic flight, being one of hypersonic aircraft technology key technology urgently to be resolved hurrily, the core of body/Propulsion Integrated is then the integration of aircraft body and air intake duct. Consider from design angle, totally the requirement of the two is also existed difference: the requirement of body is mainly high lift-drag ratio, high dischargeable capacity, and good leading edge Aerodynamic Heating barrier propterty; Requirement to air intake duct is then provide effective source of the gas as much as possible with minimum flowed energy loss for combustor. Good body-Propulsion Integrated configuration can meet designer's integration requirement to hypersonic aircraft aero-propulsive performance.

Rider design concept is applied to hypersonic aircraft body-air intake duct integrated design mainly two big advantages: one is the air-flow after can catching precommpression efficiently. This is because be possible not only to realize the purpose of precommpression air-flow by the compression of the leading edge shock of Waverider, and due to rider design, to make aerodynamic configuration lower surface higher-pressure region escape towards the air-flow of upper surface low-pressure area less, therefore can catch air-flow as much as possible. Two is by optimizing design (such as choosing suitable Angle of Shock Waves), it is possible to achieve the high lift-drag ratio performance design of aircraft.

Rider design concept is applied to the conventional method of hypersonic aircraft body-air intake duct integrated design and is mainly aircraft precursor-air intake duct rider integrated design, is referred to as waverider forebody derived-air intake duct integrated design.As depicted in figs. 1 and 2, Waverider 1 is used as hypersonic aircraft precursor, referred to as waverider forebody derived 1, air intake duct adopts lower jaw formula air intake duct, and waverider forebody derived 1 is as the precommpression face of air intake duct, the air-flow after precommpression is provided for air intake duct, waverider forebody derived 1 produces leading edge shock 5, and leading edge shock 5 is incident on inlet lip 2, and produces reflected shock wave 6, air-flow enters air intake duct distance piece 4 through the compression of leading edge shock 5, reflected shock wave 6 and air intake duct outer housing 3, provides source of the gas for combustor. In this conventional Waverider-air intake duct integrated design method, only Waverider is used as aircraft precursor, does not consider the rider integrated design (such as wing) at other positions of aircraft, therefore can not give full play to the high lift-drag ratio characteristic of Waverider.

Summary of the invention

For the defect that prior art exists, it is an object of the invention to provide a kind of hypersonic aircraft precursor, air intake duct and wing rider integrated design method, it is possible to give full play to Waverider efficient capture precommpression stream condition and the big characteristic of high lift-drag ratio two.

The technical scheme is that

A kind of hypersonic aircraft precursor, air intake duct and wing rider integrated design method, adopt following steps:

S1. in design is a kind of, supersonic speed axisymmetric flow field is compressed in outer cone mixing, and as the benchmark flow field generating hypersonic aircraft precursor-air intake duct integration configuration, claiming this benchmark flow field is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field;

S1.1 gives tip revolving body bus 10-11, the rotating shaft of tip revolving body is X-axis, and the starting point of tip revolving body bus is point 10, and the distal point of tip revolving body bus is point 11, then choosing the cross section 12 at inlet lip place, described cross section is the plane perpendicular with X-axis;

Using Supersonic Stream condition 7 and tip revolving body bus 10-11 as input parameter, utilize and have rotation characteristic line method to solve the position coordinates on the characteristic curve grid node in leading edge shock 15 and leading edge shock dependence district 16 and flow parameter, wherein junction point 10 is leading edge shock 15 with the curve 10-13 of point 13, leading edge shock 15, curve 10-14 and left lateral characteristic curve 14-13 area defined and leading edge shock rely on district 16; Point 13 is the intersection point of leading edge shock 15 and the cross section 12 at inlet lip place, and point 14 is the intersection point of the left lateral characteristic curve through point 13 and tip revolving body bus 10-11;

The curve 14-13 of S1.2 junction point 14 and point 13 is left lateral characteristic curve, by the curved section 14-11 on left lateral characteristic curve 14-13 and tip revolving body bus 10-11, utilize the intersection point 17 having rotation characteristic line method to solve the right lateral characteristic curve through point 13 and tip revolving body bus 10-11, and solve by the flow field of left lateral characteristic curve 14-13, right lateral characteristic curve 13-17 and curve 14-17 institute enclosing region;

Point 13 is as the starting point of the reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, flow direction after given reflected shock wave ripple is angular distribution, utilize the alternative manner estimated-correct, solve the position of reflected shock wave and the intersection point 18 of reflected shock wave and tip revolving body bus 10-11, claiming this reflected shock wave is reflected shock wave 13-18, the flow parameter after then utilizing oblique shock wave relational expression to solve reflected shock wave 13-18 ripple; By left lateral characteristic curve 14-13, reflected shock wave 13-18 and the defined region of curve 14-18 19 as the main compressional zone of isentropic Compression between the shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field;Wherein, flow direction angle is the angle of flow direction and the axial coordinate axle X of cylindrical-coordinate system;

S1.3 utilizes rotation characteristic line method, by the flow parameter after reflected shock wave 13-18 ripple, solve the outer housing internal face leading portion curve 13-31 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, until meeting at a little 31 with the right lateral characteristic curve crossing point 18, and solve the reflected shock wave dependence district 32 in the hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field surrounded by curve 13-31, reflected shock wave 13-18 and right lateral characteristic curve 18-31;

Centrosome wall curve 33 on the right side of S1.4, set point 18 and the Mach Number Distribution on this centrosome wall curve 33, make the centrosome wall curve 33 angle of contingence in point 18 positions overlap with local flow direction angle, then given air intake port cross section 34 simultaneously; Utilization has rotation characteristic line method, by the Mach Number Distribution on centrosome wall curve 33 and this centrosome wall curve 33, outer housing internal face back segment curve 35 on the right side of solution point 31, until air intake port cross section 34, the distal point 36 of outer housing internal face back segment curve 35 is positioned on air intake port cross section 34; Simultaneously, solving the stable region 38 in the hypersonic aircraft precursor defined for curved section 18-37 by right lateral characteristic curve 18-31, outer housing internal face back segment curve 35, right lateral characteristic curve 36-37 and centrosome wall curve 33-air intake duct integral shaft symmetric reference flow field, point 37 is the intersection point of the right lateral characteristic curve through point 36 and centrosome wall curve 33; Wherein, cut angle the angle being curve near tangent with the axial coordinate axle of cylindrical-coordinate system;

Obtain outer cone mixing compression supersonic speed axisymmetric flow field in one, it can be used as the benchmark flow field generating hypersonic aircraft precursor-air intake duct integration configuration, this benchmark flow field includes leading edge shock 15, reflected shock wave 13-18, and leading edge shock relies on the main compressional zone 19 of isentropic Compression between district 16, shock wave, reflected shock wave relies on district 32 and stable region 38.

S2. design wing axis symmetric reference flow field, this benchmark flow field is used for generating wing configuration, and claiming this benchmark flow field is wing axis symmetric reference flow field;

The basis of design tip revolving body bus 43, the leading portion curve 10 '-17 of tip revolving body bus 43 ' the curve 10-17 designed with step S1.2 identical, at curve 10 '-17 ' continues to design complete tip revolving body bus 43.

The rotating shaft of tip revolving body is X-axis, the bottom transverse cross section of tip revolving body is 40, the starting point of tip revolving body bus 43 is point 10 ', the distal point of tip revolving body bus 43 is point 39, tip revolving body bus 43 is by leading portion curve 10 '-17 ' and back segment curve 17 '-39 form, wherein leading portion curve 10 '-17 ' with step S1 in be identical for designing the tip revolving body bus 10-11 leading portion curve 10-17 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, the tip revolving body obtained is rotated under the effect of zero-incidence and Supersonic Stream 7 by tip revolving body bus 43, produce attached body leading edge shock 10 '-13 '-41, wherein, shock point 13 ' is the intersection point of the cross section 12 at inlet lip place defined in leading edge shock 10 '-13 '-41 and step S1, shock wave section 10 '-13 ' with step S1 in leading edge shock section 10-13 be identical, by leading edge shock section 10 '-13 ', the defined flow field of curve 13 '-18 ' and curve 10 '-18 ' with in step S1 by leading edge shock 15, flow field defined for reflected shock wave 13-18 and curve 10-18 is also identical, wherein, the position of the point 18 in the position of point 18 ' and step S1 is identical, curve 13 '-18 ' shape and position be also identical with the shape of the reflected shock wave 13-18 in step S1 and position, the position of the point 14 in point 14 ' and step S1 is also identical.

Using Supersonic Stream condition 7 and tip revolving body bus 43 as input parameter, utilization has rotation characteristic line method, solve the supersonic speed axisymmetric flow field around the tip revolving body that zero-incidence bus is 43, obtain the position coordinates on the characteristic curve grid node after leading edge shock and shock wave ripple and flow parameter, wherein, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed and local flow direction angle, the position coordinates on characteristic curve grid node on leading edge shock 10 '-13 '-41 can represent leading edge shock profile. it is wing axis symmetric reference flow field by leading edge shock 10 '-13 '-41, tip revolving body bus 43 and straight line 41-39 area defined.

S3. given aircraft precursor costa, inlet lip molded line, air intake duct sweepforward side plate costa and the leading edge of a wing line drop shadow curve in bottom transverse cross section; From precursor costa, inlet lip molded line and air intake duct sweepforward side plate costa, hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field carries out streamlined impeller, generate hypersonic aircraft precursor-air intake duct integration configuration; From leading edge of a wing line, wing axis symmetric reference flow field carries out streamlined impeller, generate wing configuration; Hypersonic aircraft precursor-air intake duct integration configuration and wing configuration collectively constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

Further, in the step S1.1 of the present invention, the selection principle of the cross section 12 at inlet lip place is this cross section with point 10 along the distance of X-direction more than aircraft precursor length.

Further, in the step S1.1 of the present invention, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle, and the position coordinates on characteristic curve grid node on leading edge shock 15 can represent leading edge shock profile. Described characteristic curve grid node is the intersection point of left lateral characteristic curve and right lateral characteristic curve.

Further, in the step S1.2 of the present invention, utilizing the alternative manner estimated-correct to solve the position of reflected shock wave 13-18, its method is as follows:

The starting point of reflected shock wave 13-18 is point 13, the intersection point of reflected shock wave 13-18 and left lateral characteristic curve is called shock point, the position solving reflected shock wave 13-18 is to solve for the coordinate figure of all shock points, until the intersection point 18 of reflected shock wave 13-18 and tip revolving body bus 10-11.

For the shock point 22 and 23 that any two on reflected shock wave 13-18 is adjacent, it is defined as upstream shock point 22 near the shock point 22 putting 13, the shock point of point of distance 13 is defined as downstream shock point 23, the coordinate figure of upstream shock point 22 the coordinate figure method solving downstream shock point 23 is as described below:

Characteristic curve grid node is the intersection point of left lateral characteristic curve and right lateral characteristic curve, the position coordinates of characteristic curve grid node and flow parameter all can pass through have rotation characteristic line method to solve and obtain, the position coordinates of characteristic curve grid node is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, and flow parameter includes local static pressure, local density, local speed, local flow direction angle;

Shown in predicting equation in the alternative manner estimated-correct such as formula (1), shown in the iterative equation of correction such as formula (2):

$r_{i+1}^0=r_i+tan(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)\mathrm{\Δ}x---\left(1\right)$

$r_{i+1}^n=r_i+\tan \[\frac{(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)+(\mathrm{\π}-({\mathrm{\β}}_{i+1}^{n-1}-{\mathrm{\θ}}_{i+1,1}^{n-1}\left)\right)}{2}\]\mathrm{\Δ}x---\left(2\right)$

Wherein, x is the shock point coordinate at the axial coordinate axle of cylindrical-coordinate system, and r is the shock point coordinate at the radial coordinate axle of cylindrical-coordinate system, r_iValue for the radial coordinate axle in cylindrical-coordinate system of upstream shock point 22, i is the Position Number of shock point, Δ x is downstream shock point 23 and upstream shock point 22 difference in X-direction, and β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction;It is the r value after downstream shock point 23 is estimated,It is that downstream shock point 23 corrects r value obtained after n time; θ_i,1It is the local flow direction angle θ value of the wavefront of upstream shock point 22,It is the downstream shock point 23 local flow direction angle θ value that corrects wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; β_iIt is the β value of upstream shock point 22,It is that downstream shock point 23 corrects β value obtained after n-1 time,Solved by formula (3) and obtain:

$tan({\mathrm{\θ}}_{i+1,1}^{n-1}-{\mathrm{\θ}}_{i+1,2})=2\cot \mathrm{\β}\frac{{\left(M_{i+1,1}^{n-1}\right)}^2{\sin }^2{\mathrm{\β}}_{i+1}^{n-1}-1}{{\left(M_{i+1,1}^{n-1}\right)}^2(\mathrm{\γ}+cos(2{\mathrm{\β}}_{i+1}^{n-1}))+2}---\left(3\right)$

Wherein,WithRespectively downstream shock point 23 corrects local Mach number M value and the local flow direction angle θ value of wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; θ_i+1,2It is the local flow direction angle θ value after the ripple of downstream shock point 23, θ_i+1,2It is known conditions, θ_i+1,2Obtain according to the flow direction after reflected shock wave 13-18 ripple is angular distribution.

Further, in the step S1.2 of the present invention, described oblique shock wave relational expression is utilized to solve after reflected shock wave ripple the formula of flow parameter as shown in (4)～(8):

$tan\left(\mathrm{\Δ}\mathrm{\θ}\right)=2\cot \mathrm{\β}\frac{M_1^2{\sin }^2\mathrm{\β}-1}{M_1^2(\mathrm{\γ}+cos2\mathrm{\β})+2}---\left(4\right)$

Δ θ=θ₁-θ₂(5)

$\frac{P_2}{P_1}=\frac{2\mathrm{\γ}}{\mathrm{\γ}+1}(M_1^2{\sin }^2\mathrm{\β}-\frac{\mathrm{\γ}-1}{2\mathrm{\γ}})---\left(6\right)$

$\frac{{\mathrm{\ρ}}_1}{{\mathrm{\ρ}}_2}=\frac{2}{\mathrm{\γ}+1}(\frac{1}{M_1^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{2})---\left(7\right)$

$\frac{V_2}{V_1}=\frac{sin\mathrm{\β}}{sin\[\mathrm{\β}-\mathrm{\Δ}\mathrm{\θ}\]}(\frac{2}{(\mathrm{\γ}+1)M^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{\mathrm{\γ}+1})---\left(8\right)$

Wherein, β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction, and Δ θ is the local flow-deviation angle of reflected shock wave, θ₁It is the local flow direction angle of reflected shock wave wavefront, M₁It is the local Mach number of reflected shock wave wavefront, P₁It is the local static pressure of reflected shock wave wavefront, ρ₁It is the local density of reflected shock wave wavefront, V₁It is the local speed of reflected shock wave wavefront, θ₂It is the local flow direction angle after reflected shock wave ripple, P₂It is the local static pressure after reflected shock wave ripple, ρ₂It is the local density after reflected shock wave ripple, V₂It is the local speed after reflected shock wave ripple.

Further, in the step S1.2 of the present invention, the angle in the infinitesimal of the reflected shock wave at shock point (upstream or downstream shock point) place and the reflected shock wave velocity of wave front direction at shock point place is the reflected shock wave local Angle of Shock Waves β in shock point position, and the angle of the reflected shock wave velocity of wave front direction at shock point place and the axial coordinate axle of cylindrical-coordinate system is the reflected shock wave wavefront flow direction angle θ in shock point position₁, after the reflected shock wave ripple at shock point place, velocity attitude is reflected shock wave flow direction angle θ after the ripple of shock point position with the angle of the axial coordinate axle of cylindrical-coordinate system₂, behind the reflected shock wave velocity of wave front direction at shock point place and the reflected shock wave ripple at shock point place, the angle of velocity attitude is the reflected shock wave local flow-deviation angle Δ θ in shock point position.

Further, the method for the step S3 of the present invention is:

Given aircraft body costa drop shadow curve 47-48-49-50-51, this curve is as the aircraft body costa drop shadow curve in bottom transverse cross section 40, wherein, point 48 and 50 is the intersection point of aircraft body costa drop shadow curve and the shock wave contour line 45 on the cross section 12 at inlet lip place, and point 47 and point 51 are the intersection points of aircraft body costa drop shadow curve and the shock wave contour line 46 on bottom transverse cross section 40;

The center of circle of shock wave contour line 45 and shock wave contour line 46 overlaps, the left side ray 44-52 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve 47-48-49-50-51 meet at a little 52, and meet at a little 54 with shock wave contour line 45, the right side ray 44-53 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve 47-48-49-50-51 meet at a little 53, and meet at a little 55 with shock wave contour line 45, left side ray 44-52, the angle in right side ray 44-53 and the longitudinally asymmetric face 64 of aircraft is Φ value, curved section between point 52, point 49 and point 53, as the precursor costa drop shadow curve in bottom transverse cross section 40, is called precursor costa drop shadow curve 52-49-53, the curved section between curved section and point 53 and point 51 between point 47 and point 52, as the leading edge of a wing line drop shadow curve in bottom transverse cross section 40, is called leading edge of a wing line projection curve 47-52 and leading edge of a wing line projection curve 53-51, curved section between point 54 and point 55, as the inlet lip molded line drop shadow curve in bottom transverse cross section 40, is called inlet lip molded line drop shadow curve 54-55,

Application free-streamline method, by precursor costa drop shadow curve 52-49-53, inlet lip molded line drop shadow curve 54-55 and leading edge of a wing line projection curve 47-52 and leading edge of a wing line projection curve 53-51, calculate precursor costa, inlet lip molded line and leading edge of a wing line respectively;

From precursor costa and inlet lip molded line, at hypersonic aircraft precursor-air intake duct integral shaft symmetric reference, flow field carries out streamlined impeller, solve all streamlines through precursor costa and inlet lip molded line, until air intake port cross section 34 position, and then obtain air intake port molded line and closed loop curve 56-57-58-59-60-61; From leading edge of a wing line, in wing axis symmetric reference flow field, carry out streamlined impeller, solve all streamlines through leading edge of a wing line, until position, bottom transverse cross section 40, and then obtain trailing edge line 47-62 and 63-51;

All streamline setting-outs through precursor costa are become stream interface 65, all streamline setting-outs through inlet lip molded line is become stream interface 66, and forms hypersonic aircraft precursor-air intake duct integration configuration plus air intake duct sweepforward side plate 67,65,66 and 67;

All streamline setting-outs on the left of leading edge of a wing line are become stream interface 68, and all streamline setting-outs on the right side of leading edge of a wing line are become stream interface 69, upper surface application free-streamline method generates, and forms wing configuration, and described wing configuration includes port wing 70 and starboard wing 71;

Described hypersonic aircraft precursor-air intake duct integration configuration and wing configuration constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

Further, in the step S3 of the present invention, free-streamline method the implementation generating precursor costa and inlet lip molded line is as follows:

Set up an office 72 is a discrete point on precursor costa drop shadow curve 52-49-53 or inlet lip molded line drop shadow curve 54-55, intersect at a point with leading edge shock 10-13 with through point 72 straight line 73 parallel with the axial coordinate axle X of cylindrical-coordinate system, this intersection point is the point on precursor costa or inlet lip molded line, being referred to as precursor leading edge point or inlet lip point 74, straight line 74-72 is the free-stream line through precursor leading edge point or inlet lip point 74;

From precursor leading edge point or inlet lip point 74s, using the position coordinates on each characteristic curve grid node in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field and flow parameter as known conditions, streamlined impeller method is utilized to solve streamline 75, until air intake port cross section 34, the streamline 75 distal point on air intake port cross section 34 is the point on air intake port molded line, by title air intake port point 76;

Use above-mentioned same procedure, solve and obtain all precursor leading edge point and inlet lip point, and all streamlines through precursor leading edge point and inlet lip point, and obtain the air intake port point corresponding with precursor leading edge point and the air intake port point corresponding with inlet lip point; All precursor leading edge point composition precursor costa, all inlet lip points composition inlet lip molded line, the upper wall surface 56-57-58 of all air intake port point composition air intake port molded line corresponding with precursor leading edge point, the lower wall surface 61-60-59 of all air intake port point composition air intake port molded line corresponding with inlet lip point.

Further, in the step S3 of the present invention, the described implementation by free-streamline method generation leading edge of a wing line and generation trailing edge line is as follows:

Set up an office 77 is a discrete point on leading edge of a wing line projection curve, intersect at a point with leading edge shock 10 '-13 '-41 with through point 77 straight line 78 parallel with the axial coordinate axle X of cylindrical-coordinate system, this intersection point is the point on leading edge of a wing line, it is called leading edge of a wing point 79, and straight line 79-77 is the free-stream line through leading edge point 79;

From leading edge of a wing point 79s, using the position coordinates on characteristic curve grid node in wing axis symmetric reference flow field and flow parameter as known conditions, streamlined impeller method is utilized to solve streamline 80, until bottom transverse cross section 40, the streamline 80 distal point on bottom transverse cross section 40 is the point on trailing edge line, and this point is called trailing edge point 81;

Use above-mentioned same procedure, solve and obtain all leading edges of a wing point, and all streamlines through leading edge of a wing point, and obtain all trailing edges point, all port wing leading edge point and starboard wing leading edge point separately constitute port wing costa and starboard wing costa, and all port wing trailing edge points and starboard wing trailing edge point separately constitute port wing trailing edge line and starboard wing trailing edge line; Port wing costa and starboard wing costa composition leading edge of a wing line, port wing trailing edge line and starboard wing trailing edge line composition trailing edge line.

Further, in the step S3 of the present invention, the implementation of described air intake duct sweepforward side plate costa and air intake duct sweepforward side plate is as follows:

Air intake duct sweepforward side plate includes the costa of left plate and right plate, left plate and the right plate drop shadow curve in bottom transverse cross section 40 respectively curve 52-54 and curve 53-55, and side plate costa is arranged on reflected shock wave 13-18;

Intersecting at side plate leading edge point 83 with through point 52 straight line 82 parallel with the axial coordinate axle X of cylindrical-coordinate system with reflected shock wave 13-18, side plate leading edge point 83 is the distal point on left plate costa; Side plate leading edge point 85 is intersected at reflected shock wave 13-18 with through point 54 straight line 84 parallel with the axial coordinate axle X of cylindrical-coordinate system, side plate leading edge point 85 is the starting point on left plate costa, on two dimensional surface, side plate leading edge point 85 overlaps with the point 13 on reflected shock wave;

From side plate leading edge point 83s, the position coordinates reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field relied in district 32 and stable region 38 on characteristic curve grid node and flow parameter are as known conditions, streamlined impeller method is utilized to solve streamline 86, until air intake port cross section 34, the streamline 86 distal point on air intake port cross section 34 is the point on air intake duct side plate trailing edge line, and this distal point is called side plate trailing edge point 87;

By above-mentioned identical method, solve and obtain all side plate leading edge point, and all streamlines through side plate leading edge point, and obtain all side plate trailing edge points; All streamline setting-outs through left plate leading edge point are become stream interface, forms left plate, all left plate leading edge point are formed left plate costa, all left plate trailing edge points are formed the left side wall 56-61 of air intake port molded line; All streamline setting-outs through right plate leading edge point are become stream interface, forms right plate, all right plate leading edge point are formed right plate costa, all right plate trailing edge points are formed the right side wall 58-59 of air intake port molded line; Left plate costa and right plate costa composition air intake duct sweepforward side plate costa.

The invention has the beneficial effects as follows:

The present invention is on waverider forebody derived/air intake duct integrated design basis, application theory of characteristics, build hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field and Liang Ge flow field, wing axis symmetric reference flow field, then applying streamlined impeller technology in two benchmark flow fields respectively and generate hypersonic aircraft precursor-air intake duct integration configuration and wing configuration, two configurations collectively constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration. Under design point, whole hypersonic aircraft Flow Field outside has class rider characteristic, and waverider forebody derived is as the precommpression face of air intake duct, and efficient capture precommpression air-flow is supplied to air intake duct, and rider wing provides high lift-drag ratio for aircraft.

Accompanying drawing explanation

Fig. 1 is the schematic three dimensional views of conventional waverider forebody derived-air intake duct integrated design scheme;

Fig. 2 is the schematic diagram of longitudinally asymmetric of conventional waverider forebody derived-air intake duct integrated design scheme;

The leading edge shock that Fig. 3 is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field relies on district;

Fig. 4 is main compressional zone and the reflected shock wave of isentropic Compression between hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field shock wave;

Fig. 5 is the schematic diagram solving reflected shock wave position;

Fig. 6 is the local Angle of Shock Waves β of reflected shock wave, the flow direction angle θ of reflected shock wave wavefront₁, flow direction angle θ after reflected shock wave ripple₂And the definition of the local flow-deviation angle Δ θ of reflected shock wave;

The reflected shock wave that Fig. 7 is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field relies on district;

Fig. 8 is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field and stable region thereof;

Fig. 9 is wing axis symmetric reference flow field;

Figure 10 illustrates the shock wave of aircraft precursor costa, inlet lip molded line, air intake duct sweepforward side plate costa, leading edge of a wing line and cross section 12 and position, bottom transverse cross section 40 drop shadow curve in bottom transverse cross section 40;

Figure 11 is hypersonic aircraft precursor-air intake duct integration configuration and the stream interface constructing it;

Figure 12 is wing configuration and the stream interface constructing it;

Figure 13 is the hypersonic aircraft precursor-air intake duct-wing rider integration configuration being made up of hypersonic aircraft precursor-air intake duct integration configuration and wing configuration;

Figure 14 is the design diagram of precursor costa, inlet lip molded line, streamline and air intake port molded line;

Figure 15 is the design diagram of leading edge of a wing line, streamline and trailing edge line;

Figure 16 is air intake duct sweepforward side plate costa, streamline;

Figure 17 is the streamline in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field through leading edge point 88;

Figure 18 is the streamline in wing axis symmetric reference flow field through leading edge point 88.

In figure, 1 represents waverider forebody derived; 2 represent inlet lip; 3 represent air intake duct outer housing; 4 represent air intake duct distance piece; 5 represent the leading edge shock produced by waverider forebody derived; 6 represent that leading edge shock 5 is incident on the reflected shock wave of lip 2; 7 represent Supersonic Stream; 10 represent tip revolving body summit; X represents the axial coordinate axle of cylindrical-coordinate system; Y represents the radial coordinate axle of cylindrical-coordinate system; The 11 end wall cake representing tip revolving body bus 10-11; 12 represent inlet lip cross section; 13 represent the leading edge shock 15 distal point at inlet lip cross section 12 place; The 14 wall intersection points representing the left lateral characteristic curve through point 13 and tip revolving body bus 10-11; 15 represent around the tip revolving body that bus is 10-11, leading edge shock before inlet lip cross section 12; 16 represent that the leading edge shock in the hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field defined by leading edge shock 15, curve 10-14 and left lateral characteristic curve 14-13 relies on district; 17 represent by the wall intersection point of the right lateral characteristic curve with tip revolving body bus 10-11 passing through point 13; The 18 wall intersection points representing reflected shock wave and tip revolving body bus 10-11; 19 represent the main compressional zone of isentropic Compression between left lateral characteristic curve 14-13, reflected shock wave 13-18 and the defined hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field shock wave of curve 14-18; 20 and 21 represent the adjacent feature wire grid node on same left lateral characteristic curve; 22 and 23 represent the adjacent shock point on reflected shock wave 13-18; The infinitesimal of 24 reflected shock waves representing point 23 places; The 25 reflected shock wave velocity of wave front directions representing point 23 places; Velocity attitude after the 26 reflected shock wave ripples representing point 23 places; The 27 local Angle of Shock Waves β representing point 23 place's reflected shock waves; 28 represent point 23 place reflected shock wave wavefront airflow direction angle θ₁; Airflow direction angle θ after 29 expression point 23 place's reflected shock wave ripples₂; 30 represent point 23 reflected shock wave locality, place flow-deviation angle Δ θ; The 31 wall intersection points representing the right lateral characteristic curve through point 18 and hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field outer housing internal face; 32 represent that the reflected shock wave in the hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field defined by curve 13-31, reflected shock wave 13-18 and right lateral characteristic curve 18-31 relies on district; 33 represent the centrosome wall curve on the right side of point 18; 34 represent air intake port cross section; 35 represent the hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field outer housing internal face curve on the right side of point 31; 36 distal points representing hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field outer housing internal face curve 35; The 37 wall intersection points representing the right lateral characteristic curve through point 36 and hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field centrosome wall curve 33; 38 represent by the stable region in hypersonic aircraft precursor defined for the curved section 18-37-air intake duct integral shaft symmetric reference flow field of right lateral characteristic curve 18-31, outer housing internal face back segment curve 35, right lateral characteristic curve 36-37 and centrosome wall curve 33;39 represent the tip revolving body bus 43 distal point in revolving body bottom transverse cross section 40; 40 represent the tip revolving body bottom transverse cross section that bus is 43; 41 represent around the leading edge shock 10-13-41 of the tip revolving body that bus is 43 distal point in bottom transverse cross section 40; 42 represent wing axis symmetric reference flow field; 43 represent the tip revolving body bus for generating wing axis symmetric reference flow field; 44 centers of circle representing shock wave contour line 45 and 46; 45 represent the leading edge shock 15 contour line in cross section 12 position, and this contour line is a circle; 46 represent the leading edge shock 10-13-41 contour line in cross section 40 position, and this contour line is a circle; 47 intersection points representing leading edge of a wing line projection curve 47-52 and shock wave contour line 46; 48 intersection points representing leading edge of a wing line projection curve 47-52 and shock wave contour line 45; 50 intersection points representing leading edge of a wing line projection curve 53-51 and shock wave contour line 45; 51 intersection points representing leading edge of a wing line projection curve 53-51 and shock wave contour line 46; 52 left end points representing precursor costa drop shadow curve 52-53; 53 right endpoints representing precursor costa drop shadow curve 52-53; Φ is ray 44-52 or 44-53 sent by the center of circle 44 and the angle in the longitudinally asymmetric face 64 of aircraft, referred to as half exhibition angle Φ value; 54 and 55 is the left and right end points of inlet lip molded line drop shadow curve 54-55; The left side end points of the 56 upper wall surface 56-57-58 representing air intake port molded line (closed loop curve 56-57-58-59-60-61), it is the air intake port point corresponding with point 52; The midpoint of the 57 upper wall surface 56-57-58 representing air intake port molded line (closed loop curve 56-57-58-59-60-61), it is the air intake port point corresponding with point 49; The right side end points of the 58 upper wall surface 56-57-58 representing air intake port molded line (closed loop curve 56-57-58-59-60-61), it is the air intake port point corresponding with point 53; The right side end points of the 59 lower wall surface 61-60-59 representing air intake port molded line (closed loop curve 56-57-58-59-60-61), it is the air intake port point corresponding with point 55; The midpoint of the 60 lower wall surface 61-60-59 representing air intake port molded line (closed loop curve 56-57-58-59-60-61); The left side end points of the 61 lower wall surface 61-60-59 representing air intake port molded line (closed loop curve 56-57-58-59-60-61), it is the air intake port point corresponding with point 54; 62 right endpoints representing trailing edge line 47-62, it is the trailing edge point corresponding with point 52; 63 left end points representing trailing edge line 63-51, it is the trailing edge point corresponding with point 53; 64 represent longitudinally asymmetric position of aircraft; 65 represent the stream interface become by all streamline setting-outs through precursor costa; 66 represent the stream interface become by all streamline setting-outs through inlet lip molded line; 67 represent air intake duct sweepforward side plate; 68 represent the stream interface that all streamline setting-outs on the left of leading edge of a wing line become; 69 represent the stream interface that all streamline setting-outs on the right side of leading edge of a wing line become; 70 and 71 represent left and right wing respectively; 72 represent a discrete point on precursor costa drop shadow curve (52-49-53) or inlet lip molded line drop shadow curve (54-55); 73 represent through point 72 straight line parallel with the axial coordinate axle X of cylindrical-coordinate system; 74 represent the precursor leading edge point corresponding with point 72 or inlet lip point, are the intersection point of straight line 73 and leading edge shock 15 (i.e. 10-13);75 represent the streamline in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field through point 74; 76 represent the streamline 75 distal point on air intake port cross section 34, and it is also the point on air intake port molded line, i.e. air intake port point; 77 represent a discrete point in leading edge of a wing line projection curve (47-52,53-51); 78 represent through point 77 straight line parallel with the axial coordinate axle X of cylindrical-coordinate system; 79 represent the leading edge of a wing point corresponding with point 77, are the intersection point of straight line 78 and leading edge shock 10 '-13 '-41; 80 represent the streamline in wing axis symmetric reference flow field through point 79; 81 represent the streamline 80 distal point on bottom transverse cross section 40, and it is also the point on trailing edge line, i.e. trailing edge point; 82 represent through point 52 straight line parallel with the axial coordinate axle X of cylindrical-coordinate system; 83 represent the distal point on left plate costa, are the intersection point of straight line 82 and reflected shock wave 13-18; 84 represent through point 54 straight line parallel with the axial coordinate axle X of cylindrical-coordinate system; 85 represent the starting point on left plate costa, are the intersection point of straight line 84 and reflected shock wave 13-18, and on two dimensional surface, intersection point 85 overlaps with shock point 13; 88 is some leading edge point corresponding to 52, and it is the right endpoint of the left end point of precursor costa and port wing costa; 89 is the streamline in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field through leading edge point 88; 90 is the intersection point of streamline 89 and reflected shock wave 13-18; 91 is the intersection point of streamline 89 and cross section 34; 92 is the streamline in wing axis symmetric reference flow field through leading edge point 88; 93 is streamline 92 and 13 '-18 ' intersection point; 94 is the intersection point of streamline 92 and cross section 40; 10 ', 14 ', 17 ' and 18 ' is the point on tip revolving body bus 43, their definition and position and identical with position with the definition of 18 with 10,14,17 respectively; 13 ' is the intersection point of the cross section 12 at inlet lip place defined in leading edge shock 10 '-13 '-41 and Fig. 3.

Detailed description of the invention

Below in conjunction with accompanying drawing, a kind of hypersonic aircraft precursor provided by the invention, air intake duct and wing rider integrated design method are described in detail.

Step S1, application theory of characteristics, outer cone mixing compression supersonic speed axisymmetric flow field in design is a kind of, as the benchmark flow field generating hypersonic aircraft precursor-air intake duct integration configuration, claiming this benchmark flow field is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field.

S1.1, as shown in Figure 3, given tip revolving body bus 10-11, the rotating shaft of tip revolving body is X-axis, the starting point of tip revolving body bus is point 10, the distal point of tip revolving body bus is point 11, choosing the cross section 12 at inlet lip place, the selection principle of the cross section 12 at inlet lip place is this cross section with point 10 along the distance of X-direction more than aircraft precursor length. Described cross section is the plane perpendicular with X-axis.

Using Supersonic Stream condition 7 and tip revolving body bus 10-11 as input parameter, Supersonic Stream condition includes incoming flow Mach, incoming flow static pressure and incoming flow static temperature, utilization has rotation characteristic line method, and (having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ") solve the position coordinates on the characteristic curve grid node in leading edge shock 15 and leading edge shock dependence district 16 and flow parameter, wherein curve 10-13 is leading edge shock 15, by leading edge shock 15, curve 10-14 and left lateral characteristic curve 14-13 area defined and leading edge shock rely on district 16,Point 13 is the intersection point of leading edge shock 15 and the cross section 12 at inlet lip place, and point 14 is the intersection point of the left lateral characteristic curve through point 13 and tip revolving body bus 10-11. Position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle, and the position coordinates on characteristic curve grid node on leading edge shock 15 can represent leading edge shock profile. Described characteristic curve grid node is the intersection point of left lateral characteristic curve and right lateral characteristic curve.

S1.2, as shown in Figure 4, utilization has rotation characteristic line method, having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", curve 14-13 between point 14 and point 13 is left lateral characteristic curve, by the curved section 14-11 on left lateral characteristic curve 14-13 and tip revolving body bus 10-11, solve the intersection point 17 of the right lateral characteristic curve through point 13 and tip revolving body bus 10-11, and solve by left lateral characteristic curve 14-13, the flow field of right lateral characteristic curve 13-17 and curve 14-17 institute enclosing region, point 13 is as the starting point of the reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, flow direction after given reflected shock wave ripple is angular distribution, utilize the alternative manner estimated-correct, solve the position of reflected shock wave and the intersection point 18 of reflected shock wave and tip revolving body bus 10-11, claiming this reflected shock wave is reflected shock wave 13-18, then utilizes oblique shock wave relational expression to solve flow parameter after reflected shock wave 13-18 ripple. by left lateral characteristic curve 14-13, reflected shock wave 13-18 and the defined region of curve 14-18 19 as the main compressional zone of isentropic Compression between the shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field. wherein, flow direction angle is the angle of flow direction and the axial coordinate axle X of cylindrical-coordinate system.

The position concrete grammar that the alternative manner that described utilization is estimated-corrected solves reflected shock wave 13-18 is as follows.

As shown in Figure 5, fine line in Fig. 5 represents left lateral characteristic curve, dotted line represents right lateral characteristic curve, hollow node on behalf characteristic curve grid node, the starting point of reflected shock wave 13-18 is that the intersection point of point 13, reflected shock wave 13-18 and left lateral characteristic curve is referred to as shock point, for instance upstream shock point 22, downstream shock point 23, the described position solving reflected shock wave 13-18 is to solve for the coordinate figure of all shock points, until the intersection point 18 of reflected shock wave 13-18 and tip revolving body bus 10-11.

For the shock point 22 and 23 that any two on reflected shock wave 13-18 is adjacent, it is defined as upstream shock point 22 near the shock point 22 putting 13, the shock point 23 of point of distance 13 is defined as downstream shock point 23, the coordinate figure of upstream shock point 22 the coordinate figure method solving downstream shock point 23 is as described below.

In Fig. 5, the position coordinates of characteristic curve grid node and flow parameter all can pass through have rotation characteristic line method to solve and obtain, having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics "; the left Crow of M.J.; J.D. Huffman; National Defense Industry Press; 1984; p138-195 ", position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, and flow parameter includes local static pressure, local density, local speed, local flow direction angle.

Shown in described predicting equation such as formula (1), shown in the iterative equation of correction such as formula (2).

$r_{i+1}^0=r+tan(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)\mathrm{\Δ}x---\left(1\right)$

$r_{i+1}^n=r_i+tan\[\frac{(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)+(\mathrm{\π}-({\mathrm{\β}}_{i+1}^{n-1}-{\mathrm{\θ}}_{i+1,1}^{n-1}\left)\right)}{2}\]\mathrm{\Δ}x---\left(2\right)$

Wherein, x is the coordinate of the axial coordinate axle of cylindrical-coordinate system, and r is the coordinate of the radial coordinate axle of cylindrical-coordinate system, r_iValue for the radial coordinate axle in cylindrical-coordinate system of upstream shock point 22, i is the Position Number of shock point, Δ x is downstream and the upstream shock point difference in X-direction, and β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction;It is the r value after downstream shock point 23 is estimated,It is that downstream shock point 23 corrects r value obtained after n time; θ_i,1It is the local flow direction angle θ value of the wavefront of upstream shock point 22,It is the downstream shock point 23 local flow direction angle θ value that corrects wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; β_iIt is the β value of upstream shock point 22,It is that downstream shock point 23 corrects β value obtained after n-1 time,Solved by formula (3) and obtain.

$tan({\mathrm{\θ}}_{i+1,1}^{n-1}-{\mathrm{\θ}}_{i+1,2})=2\cot \mathrm{\β}\frac{{\left(M_{i+1,1}^{n-1}\right)}^2{\sin }^2{\mathrm{\β}}_{i+1}^{n-1}-1}{{\left(M_{i+1,1}^{n-1}\right)}^2(\mathrm{\γ}+cos(2{\mathrm{\β}}_{i+1}^{n-1}))+2}---\left(3\right)$

Wherein,WithRespectively downstream shock point 23 corrects local Mach number M value and the local flow direction angle θ value of wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; θ_i+1,2It is the local flow direction angle θ value after the ripple of downstream shock point 23, θ_i+1,2Being known conditions, it can obtain according to the flow direction after reflected shock wave 13-18 ripple is angular distribution.

Described oblique shock wave relational expression is utilized to solve after reflected shock wave ripple the formula of flow parameter as shown in (4)～(8).

$tan\left(\mathrm{\Δ}\mathrm{\θ}\right)=2\cot \mathrm{\β}\frac{M_1^2{\sin }^2\mathrm{\β}-1}{M_1^2(\mathrm{\γ}+cos2\mathrm{\β})+2}---\left(4\right)$

Δ θ=θ₁-θ₂(5)

$\frac{P_2}{P_1}=\frac{2\mathrm{\γ}}{\mathrm{\γ}+1}(M_1^2{\sin }^2\mathrm{\β}-\frac{\mathrm{\γ}-1}{2\mathrm{\γ}})---\left(6\right)$

$\frac{{\mathrm{\ρ}}_1}{{\mathrm{\ρ}}_2}=\frac{2}{\mathrm{\γ}+1}(\frac{1}{M_1^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{2})---\left(7\right)$

$\frac{V_2}{V_1}=\frac{sin\mathrm{\β}}{sin\[\mathrm{\β}-\mathrm{\Δ}\mathrm{\θ}\]}(\frac{2}{(\mathrm{\γ}+1)M^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{\mathrm{\γ}+1})---\left(8\right)$

Wherein, β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction, and Δ θ is the local flow-deviation angle of reflected shock wave, θ₁It is the local flow direction angle of reflected shock wave wavefront, M₁It is the local Mach number of reflected shock wave wavefront, P₁It is the local static pressure of reflected shock wave wavefront, ρ₁It is the local density of reflected shock wave wavefront, V₁It is the local speed of reflected shock wave wavefront, θ₂It is the local flow direction angle after reflected shock wave ripple, P₂It is the local static pressure after reflected shock wave ripple, ρ₂It is the local density after reflected shock wave ripple, V₂It is the local speed after reflected shock wave ripple.

The local Angle of Shock Waves β of described reflected shock wave, the flow direction angle θ of reflected shock wave wavefront₁, flow direction angle θ after reflected shock wave ripple₂And the definition of the local flow-deviation angle Δ θ of reflected shock wave is as shown in Figure 6, it is for downstream shock point 23 in figure 6, the infinitesimal 24 of the reflected shock wave at downstream shock point 23 place is the reflected shock wave local Angle of Shock Waves β in downstream shock point 23 position with the angle 27 in the reflected shock wave velocity of wave front direction 25 at downstream shock point 23 place, and the reflected shock wave velocity of wave front direction 25 at downstream shock point 23 place is the reflected shock wave wavefront flow direction angle θ in downstream shock point 23 position with the angle 28 of the axial coordinate axle of cylindrical-coordinate system₁, after the reflected shock wave ripple at downstream shock point 23 place, velocity attitude 26 is reflected shock wave flow direction angle θ after the ripple of downstream shock point 23 position with the angle 29 of the axial coordinate axle of cylindrical-coordinate system₂, after the reflected shock wave velocity of wave front direction 25 at downstream shock point 23 place and the reflected shock wave ripple at downstream shock point 23 place, the angle 30 of velocity attitude 26 is the reflected shock wave local flow-deviation angle Δ θ in downstream shock point 23 position.

S1.3, as shown in Figure 7, utilization has rotation characteristic line method, having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", by the flow parameter after reflected shock wave 13-18 ripple, solve the outer housing internal face leading portion curve 13-31 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, until meeting at a little 31 with the right lateral characteristic curve crossing point 18, and solve by curve 13-31, the reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field that reflected shock wave 13-18 and right lateral characteristic curve 18-31 surrounds relies on district 32.

S1.4, as shown in Figure 8, first the centrosome wall curve 33 on the right side of set point 18 and the Mach Number Distribution on this curve 33, make the centrosome wall curve 33 angle of contingence in point 18 positions overlap with local flow direction angle (i.e. reflected shock wave 13-18 flow direction angle after the ripple of point 18) simultaneously, to ensure that reflected shock wave 13-18 is at point 18 position areflexias, then given air intake port cross section 34, utilization has rotation characteristic line method, and (having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ") by the Mach Number Distribution on centrosome wall curve 33 and this curve, outer housing internal face back segment curve 35 on the right side of solution point 31, until air intake port cross section 34, the distal point 36 of outer housing internal face back segment curve 35 is positioned on air intake port cross section 34, meanwhile, solve by right lateral characteristic curve 18-31, outer housing internal face back segment curve 35, the stable region 38 in the hypersonic aircraft precursor defined for curved section 18-37 of right lateral characteristic curve 36-37 and centrosome wall curve 33-air intake duct integral shaft symmetric reference flow field, point 37 is the intersection point of the right lateral characteristic curve through point 36 and centrosome wall curve 33. wherein, cut angle the angle being curve near tangent with the axial coordinate axle of cylindrical-coordinate system.

Benchmark flow field as shown in Figure 8, is outer cone mixing compression supersonic speed axisymmetric flow field in one, and the present invention names this benchmark flow field for " hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field ". The structure in this benchmark flow field may be summarized to be " Liang Bo tetra-district ". Wherein, " two ripples " includes leading edge shock 15 and reflected shock wave 13-18,4th district include leading edge shock and rely on the main compressional zone 19 of isentropic Compression between district 16, shock wave, reflected shock wave dependence district 32 and stable region 38, wherein the first two district is external compression using curved cone flow field (i.e. external compression flow field), 3rd district is contract using curved cone flow field (i.e. contract flow field), and the 4th district is used for adjusting airflow direction.

Step S2, application theory of characteristics, design and hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field have the wing axis symmetric reference flow field of identical tip revolving body bus leading portion, this benchmark flow field is used for generating wing configuration, and claiming this benchmark flow field is wing axis symmetric reference flow field.

The basis of ' the curve 10-17 designed with step S1 identical, at curve 10 '-17 ' continues to design complete tip revolving body bus 43 as it is shown in figure 9, design tip revolving body bus 43, the leading portion curve 10 '-17 of tip revolving body bus 43. the rotating shaft of tip revolving body is X-axis, tip revolving body bus is the tip revolving body bottom transverse cross section of 43 is 40, the starting point of tip revolving body bus 43 is point 10 ', the distal point of tip revolving body bus 43 is point 39, tip revolving body bus 43 is by leading portion curve 10 '-17 ' and back segment curve 17 '-39 form, wherein leading portion curve 10 '-17 ' with as shown in Figure 8 be identical for designing the tip revolving body bus 10-11 leading portion curve 10-17 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, point 10 ' is identical with the position of point 10 and point 17 respectively with the position of point 17 '. rotated the tip revolving body obtained under the effect of zero-incidence and Supersonic Stream 7 by tip revolving body bus 43 (i.e. 10 '-17 '-39), produce attached body leading edge shock 10 '-13 '-41.Wherein, shock point 13 ' is the intersection point of the cross section 12 at the inlet lip place that (in Fig. 3) is defined in leading edge shock 10 '-13 '-41 and step S1, shock wave section 10 '-13 ' it is identical with the leading edge shock section 10-13 of (shown in Fig. 8) in step S1, meanwhile, ', curve 13 '-18 ' and curve 10 '-18 by leading edge shock section 10 '-13 ' defined flow field is also identical with the flow field defined by leading edge shock 15, reflected shock wave 13-18 and curve 10-18 of (shown in Fig. 8) in step S1. Wherein, in position and the step S1 of point 18 ', the position of the point 18 of (shown in Fig. 8) is identical, curve 13 '-18 ' shape and position be also identical with the shape of reflected shock wave 13-18 (shown in Fig. 8) in step S1 and position. Point 14 ' in Fig. 9 is also identical with the position of the point 14 of (shown in Fig. 8) in step S1.

Using Supersonic Stream condition 7 and tip revolving body bus 43 (i.e. 10 '-17 '-39) as input parameter, inlet flow conditions includes incoming flow Mach, incoming flow static pressure, incoming flow static temperature, utilization has rotation characteristic line method, having rotation characteristic line method is techniques known, specifically can referring to " " aerodynamics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", solve around zero-incidence tip revolving body that (bus is 43, i.e. 10 '-17 '-39) supersonic speed axisymmetric flow field, obtain the position coordinates on the characteristic curve grid node after leading edge shock and shock wave ripple and flow parameter, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle, the position coordinates on characteristic curve grid node on leading edge shock 10 '-13 '-41 can represent leading edge shock profile. it is wing axis symmetric reference flow field by leading edge shock 10 '-13 '-41, tip revolving body bus 43 and straight line 41-39 area defined.

Step S3, given aircraft precursor costa, inlet lip molded line, air intake duct sweepforward side plate costa and the leading edge of a wing line drop shadow curve in bottom transverse cross section; From precursor costa, inlet lip molded line and air intake duct sweepforward side plate costa, hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field carries out streamlined impeller, generate hypersonic aircraft precursor-air intake duct integration configuration; From leading edge of a wing line, wing axis symmetric reference flow field carries out streamlined impeller, generate wing configuration; Hypersonic aircraft precursor-air intake duct integration configuration and wing configuration collectively constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

As shown in Figure 10, given two dimension open loop curve 47-48-49-50-51, this curve is as the aircraft body costa drop shadow curve in bottom transverse cross section 40, referred to as aircraft body costa drop shadow curve (47-48-49-50-51), wherein, point 48 and 50 is aircraft body costa drop shadow curve and the cross section 12 at inlet lip place intersection point on shock wave contour line 45, and point 47 and point 51 are aircraft body costa drop shadow curve and the bottom transverse cross section 40 intersection point on shock wave contour line 46. the center of circle of shock wave contour line 45 and shock wave contour line 46 overlaps. the left side ray 44-52 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve (47-48-49-50-51) meet at a little 52, and meet at a little 54 with shock wave contour line 45, the right side ray 44-53 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve (47-48-49-50-51) meet at a little 53, and meet at a little 55 with shock wave contour line 45, left side ray 44-52, the angle in right side ray 44-53 and the longitudinally asymmetric face 64 of aircraft is Φ value, referred to as half exhibition angle Φ value,Curved section 52-49-53 between point 52, point 49 and point 53 is as the precursor costa drop shadow curve in bottom transverse cross section 40, referred to as precursor costa drop shadow curve (52-49-53); The curved section 53-51 between curved section 47-52 and point 53 and point 51 between point 47 and point 52 is as the leading edge of a wing line drop shadow curve in bottom transverse cross section 40, referred to as leading edge of a wing line projection curve (47-52,53-51); Curved section 54-55 between point 54 and point 55 is as the inlet lip molded line drop shadow curve in bottom transverse cross section 40, referred to as inlet lip molded line drop shadow curve (54-55).

Application free-streamline method, by precursor costa drop shadow curve (52-49-53), inlet lip molded line drop shadow curve (54-55) and leading edge of a wing line projection curve (47-52,53-51), precursor costa, inlet lip molded line and leading edge of a wing line are calculated respectively.

From precursor costa and inlet lip molded line, at hypersonic aircraft precursor-air intake duct integral shaft symmetric reference, flow field carries out streamlined impeller, solve all streamlines through precursor costa and inlet lip molded line, until air intake port cross section 34 position, and then obtain air intake port molded line (closed loop curve 56-57-58-59-60-61); From leading edge of a wing line, in wing axis symmetric reference flow field, carry out streamlined impeller, solve all streamlines through leading edge of a wing line, until position, bottom transverse cross section 40, and then obtain trailing edge line (47-62,63-51).

As shown in figure 11, all streamline setting-outs through precursor costa are become stream interface 65, all streamline setting-outs through inlet lip molded line are become stream interface 66, and forms hypersonic aircraft precursor-air intake duct integration configuration plus air intake duct sweepforward side plate 67,65,66 and 67.

As shown in figure 12, all streamline setting-outs on the left of leading edge of a wing line becoming stream interface 68, and all streamline setting-outs on the right side of leading edge of a wing line are become stream interface 69, upper surface application free-streamline method generates, composition wing configuration, described wing configuration includes port wing 70 and starboard wing 71.

As shown in figure 13, described hypersonic aircraft precursor-air intake duct integration configuration (being formed by 65,66 and 67) and wing configuration (being made up of 70 and 71) constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

The described implementation by free-streamline method generation precursor costa and inlet lip molded line is as described below.

As shown in figure 14, set up an office 72 is a discrete point on precursor costa drop shadow curve (52-49-53) or inlet lip molded line drop shadow curve (54-55), intersection point 74 is intersected at leading edge shock 10-13 with through point 72 straight line 73 parallel with the axial coordinate axle X of cylindrical-coordinate system, intersection point 74 is the point on precursor costa or inlet lip molded line, referred to as precursor leading edge point or inlet lip point 74, straight line 74-72 is the free-stream line through precursor leading edge point or inlet lip point 74, from precursor leading edge point or inlet lip point 74s, using the position coordinates on each characteristic curve grid node in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field and flow parameter as known conditions (wherein: position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle), (streamlined impeller method is techniques known to utilize streamlined impeller method, specifically can referring to " " turning to Design of Inlet technique study in the streamlined impeller of feature based lineation opinion ", Wei Feng, National University of Defense Technology's academic dissertation, 2012, p67-69 ") solve streamline 75, until air intake port cross section 34, the streamline 75 distal point 76 on air intake port cross section 34 is the point on air intake port molded line, it is called for short air intake port point 76,

Use above-mentioned same procedure, solve and obtain all precursor leading edge point and inlet lip point, and all streamlines through precursor leading edge point and inlet lip point, and obtain the air intake port point corresponding with precursor leading edge point and the air intake port point corresponding with inlet lip point; All precursor leading edge point composition precursor costa, all inlet lip points composition inlet lip molded line, the upper wall surface 56-57-58 of all air intake port point composition air intake port molded line corresponding with precursor leading edge point, the lower wall surface 61-60-59 of all air intake port point composition air intake port molded line corresponding with inlet lip point.

The described implementation being generated leading edge of a wing line by free-streamline method, and the implementation generating trailing edge line is as described below:

As shown in figure 15, set up an office 77 is leading edge of a wing line projection curve (47-52, a discrete point on 53-51), intersection point 79 is intersected at leading edge shock 10 '-13 '-41 with through point 77 straight line 78 parallel with the axial coordinate axle X of cylindrical-coordinate system, intersection point 79 is the point on leading edge of a wing line, referred to as leading edge of a wing point 79, straight line 79-77 is the free-stream line through leading edge point 79, from leading edge of a wing point 79s, using the position coordinates on characteristic curve grid node in wing axis symmetric reference flow field and flow parameter as known conditions (wherein, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle), (streamlined impeller method is techniques known to utilize streamlined impeller method, specifically can referring to " " turning to Design of Inlet technique study in the streamlined impeller of feature based lineation opinion ", Wei Feng, National University of Defense Technology's academic dissertation, 2012, p67-69 ") solve streamline 80, until bottom transverse cross section 40, the streamline 80 distal point 81 on bottom transverse cross section 40 is the point on trailing edge line, referred to as trailing edge point 81,

Use above-mentioned same procedure, solve and obtain all leading edges of a wing point, and all streamlines through leading edge of a wing point, and obtain all trailing edges point, all port wing leading edge point and starboard wing leading edge point separately constitute port wing costa and starboard wing costa, and all port wing trailing edge points and starboard wing trailing edge point separately constitute port wing trailing edge line and starboard wing trailing edge line. Port wing costa and starboard wing costa composition leading edge of a wing line, port wing trailing edge line and starboard wing trailing edge line composition trailing edge line.

The implementation of described air intake duct sweepforward side plate costa and air intake duct sweepforward side plate is as described below:

Air intake duct sweepforward side plate includes left plate and right plate. As shown in figure 16, the costa of left plate and the right plate drop shadow curve in bottom transverse cross section 40 respectively curve 52-54 (two-end-point of curve 52-54 is respectively put 52 and puts 54) and 53-55 (two-end-point of curve 53-55 is respectively put 53 and puts 55), in order to make side plate be sweepforward type, side plate costa is arranged on reflected shock wave 13-18. Intersecting at side plate leading edge point 83 with through point 52 straight line 82 parallel with the axial coordinate axle X of cylindrical-coordinate system with reflected shock wave 13-18, side plate leading edge point 83 is the distal point on left plate costa; Side plate leading edge point 85 is intersected at reflected shock wave 13-18 with through point 54 straight line 84 parallel with the axial coordinate axle X of cylindrical-coordinate system, side plate leading edge point 85 is the starting point on left plate costa, on two dimensional surface, side plate leading edge point 85 overlaps with shock point 13.From side plate leading edge point 83s, the position coordinates relied on by the reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field in district 32 and stable region 38 on characteristic curve grid node and flow parameter are as known conditions (wherein, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle), (streamlined impeller method is techniques known to utilize streamlined impeller method, specifically can referring to " " turning to Design of Inlet technique study in the streamlined impeller of feature based lineation opinion ", Wei Feng, National University of Defense Technology's academic dissertation, 2012, p67-69 ") solve streamline 86, until air intake port cross section 34, the streamline 86 distal point 87 on air intake port cross section 34 is the point on air intake duct side plate trailing edge line, it is called for short side plate trailing edge point 87.

By above-mentioned identical method, solve and obtain all side plate leading edge point, and all streamlines through side plate leading edge point, and obtain all side plate trailing edge points. All streamline setting-outs through left plate leading edge point are become stream interface, forms left plate, all left plate leading edge point are formed left plate costa, all left plate trailing edge points are formed the left side wall 56-61 of air intake port molded line; All streamline setting-outs through right plate leading edge point are become stream interface, forms right plate, all right plate leading edge point are formed right plate costa, all right plate trailing edge points are formed the right side wall 58-59 of air intake port molded line; Left plate costa and right plate costa composition air intake duct sweepforward side plate costa.

The advantage of described air intake duct sweepforward side plate is when air intake duct works under low mach, by both sides overflow, reduce minimum work Mach number, particularly when air intake duct is inoperative, intrinsic pressure section of import department separates stream and can overflow from both sides, being beneficial to intake duct starting performance, namely air intake duct sweepforward side plate has automatic overflow and is beneficial to the advantage of intake duct starting.

The upper wall surface 56-57-58 of described air intake port molded line, lower wall surface 61-60-59, left side wall 56-61 and right side wall 58-59 form air intake port molded line (closed loop curve 56-57-58-59-60-61).

Described hypersonic aircraft precursor-air intake duct integration configuration is as described below with the reason that wing configuration can be integrated.

As shown in Figure 17 and Figure 18, point 88 is some leading edge point corresponding to 52, and it is the right endpoint of the left end point of precursor costa and port wing costa; In hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, the streamline through leading edge point 88 is 89, and streamline 89 and reflected shock wave 13-18 and cross section 34 meet at a little 90 and put 91 respectively, thus streamline 89 is divided into two sections of 88-90 and 90-91; In wing axis symmetric reference flow field, the streamline through leading edge point 88 is 92, streamline 92 and 13 '-18 ' and cross section 40 meet at respectively a little 93 and point 94, thus streamline 92 is divided into two sections of 88-93 and 93-94. ', curve 13 '-18 ' and curve 10 '-18 due to as described in step S2 by leading edge shock section 10 '-13 ' defined flow field be identical by the flow field that leading edge shock 15, reflected shock wave 13-18 and curve 10-18 are defined, therefore the leading portion streamline 88-90 of the streamline 89 and leading portion streamline 88-93 of streamline 92 is identical. And streamline 89 is the streamline of precursor left side edge, wherein leading portion streamline 88-90 is external compression section, streamline 92 is the streamline of port wing right side edge, thus can ensure that the external compression section of precursor left side edge streamline is identical with the leading portion of port wing right side edge streamline, so that hypersonic aircraft precursor-air intake duct integration configuration and port wing configuration are smoothly connected in edge butt joint place, namely both are integrated.Similar, hypersonic aircraft precursor-air intake duct integration configuration and starboard wing configuration are smoothly connected in edge butt joint place, and namely both are integrated.

In sum; although the present invention is disclosed above with preferred embodiment; so it is not limited to the present invention; any those of ordinary skill in the art; without departing from the spirit and scope of the present invention; when doing various change and retouching, therefore protection scope of the present invention ought be as the criterion depending on the scope that claims define.

Claims (10)

1. a hypersonic aircraft precursor, air intake duct and wing rider integrated design method, it is characterised in that comprise the following steps:

S1. in design is a kind of, supersonic speed axisymmetric flow field is compressed in outer cone mixing, and as the benchmark flow field generating hypersonic aircraft precursor-air intake duct integration configuration, claiming this benchmark flow field is hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field;

S1.1 gives tip revolving body bus 10-11, the rotating shaft of tip revolving body is X-axis, and the starting point of tip revolving body bus is point 10, and the distal point of tip revolving body bus is point 11, then choosing the cross section 12 at inlet lip place, described cross section is the plane perpendicular with X-axis;

Using Supersonic Stream condition 7 and tip revolving body bus 10-11 as input parameter, utilize and have rotation characteristic line method to solve the position coordinates on the characteristic curve grid node in leading edge shock 15 and leading edge shock dependence district 16 and flow parameter, wherein junction point 10 is leading edge shock 15 with the curve 10-13 of point 13, leading edge shock 15, curve 10-14 and left lateral characteristic curve 14-13 area defined and leading edge shock rely on district 16; Point 13 is the intersection point of leading edge shock 15 and the cross section 12 at inlet lip place, and point 14 is the intersection point of the left lateral characteristic curve through point 13 and tip revolving body bus 10-11;

The curve 14-13 of S1.2 junction point 14 and point 13 is left lateral characteristic curve, by the curved section 14-11 on left lateral characteristic curve 14-13 and tip revolving body bus 10-11, utilize the intersection point 17 having rotation characteristic line method to solve the right lateral characteristic curve through point 13 and tip revolving body bus 10-11, and solve by the flow field of left lateral characteristic curve 14-13, right lateral characteristic curve 13-17 and curve 14-17 institute enclosing region;

Point 13 is as the starting point of the reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, flow direction after given reflected shock wave ripple is angular distribution, utilize the alternative manner estimated-correct, solve the position of reflected shock wave and the intersection point 18 of reflected shock wave and tip revolving body bus 10-11, claiming this reflected shock wave is reflected shock wave 13-18, the flow parameter after then utilizing oblique shock wave relational expression to solve reflected shock wave 13-18 ripple; By left lateral characteristic curve 14-13, reflected shock wave 13-18 and the defined region of curve 14-18 19 as the main compressional zone of isentropic Compression between the shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field; Wherein, flow direction angle is the angle of flow direction and the axial coordinate axle X of cylindrical-coordinate system;

S1.3 utilizes rotation characteristic line method, by the flow parameter after reflected shock wave 13-18 ripple, solve the outer housing internal face leading portion curve 13-31 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, until meeting at a little 31 with the right lateral characteristic curve crossing point 18, and solve the reflected shock wave dependence district 32 in the hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field surrounded by curve 13-31, reflected shock wave 13-18 and right lateral characteristic curve 18-31;

Centrosome wall curve 33 on the right side of S1.4, set point 18 and the Mach Number Distribution on this centrosome wall curve 33, make the centrosome wall curve 33 angle of contingence in point 18 positions overlap with local flow direction angle, then given air intake port cross section 34 simultaneously; Utilization has rotation characteristic line method, by the Mach Number Distribution on centrosome wall curve 33 and this centrosome wall curve 33, outer housing internal face back segment curve 35 on the right side of solution point 31, until air intake port cross section 34, the distal point 36 of outer housing internal face back segment curve 35 is positioned on air intake port cross section 34; Simultaneously, solving the stable region 38 in the hypersonic aircraft precursor defined for curved section 18-37 by right lateral characteristic curve 18-31, outer housing internal face back segment curve 35, right lateral characteristic curve 36-37 and centrosome wall curve 33-air intake duct integral shaft symmetric reference flow field, point 37 is the intersection point of the right lateral characteristic curve through point 36 and centrosome wall curve 33; Wherein, cut angle the angle being curve near tangent with the axial coordinate axle of cylindrical-coordinate system;

Obtain outer cone mixing compression supersonic speed axisymmetric flow field in one, it can be used as the benchmark flow field generating hypersonic aircraft precursor-air intake duct integration configuration, this benchmark flow field includes leading edge shock 15, reflected shock wave 13-18, and leading edge shock relies on the main compressional zone 19 of isentropic Compression between district 16, shock wave, reflected shock wave relies on district 32 and stable region 38;

S2. design wing axis symmetric reference flow field, this benchmark flow field is used for generating wing configuration, and claiming this benchmark flow field is wing axis symmetric reference flow field;

The basis of design tip revolving body bus 43, the leading portion curve 10 '-17 of tip revolving body bus 43 ' the curve 10-17 designed with step S1.2 identical, at curve 10 '-17 ' continues to design complete tip revolving body bus 43;

The rotating shaft of tip revolving body is X-axis, the bottom transverse cross section of tip revolving body is 40, the starting point of tip revolving body bus 43 is point 10 ', the distal point of tip revolving body bus 43 is point 39, tip revolving body bus 43 is by leading portion curve 10 '-17 ' and back segment curve 17 '-39 form, wherein leading portion curve 10 '-17 ' with step S1 in be identical for designing the tip revolving body bus 10-11 leading portion curve 10-17 in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field, the tip revolving body obtained is rotated under the effect of zero-incidence and Supersonic Stream 7 by tip revolving body bus 43, produce attached body leading edge shock 10 '-13 '-41, wherein, shock point 13 ' is the intersection point of the cross section 12 at inlet lip place defined in leading edge shock 10 '-13 '-41 and step S1, shock wave section 10 '-13 ' with step S1 in leading edge shock section 10-13 be identical, by leading edge shock section 10 '-13 ', the defined flow field of curve 13 '-18 ' and curve 10 '-18 ' with in step S1 by leading edge shock 15, flow field defined for reflected shock wave 13-18 and curve 10-18 is also identical, wherein, the position of the point 18 in the position of point 18 ' and step S1 is identical, curve 13 '-18 ' shape and position be also identical with the shape of the reflected shock wave 13-18 in step S1 and position, the position of the point 14 in point 14 ' and step S1 is also identical,

Using Supersonic Stream condition 7 and tip revolving body bus 43 as input parameter, utilization has rotation characteristic line method, solve the supersonic speed axisymmetric flow field around the tip revolving body that zero-incidence bus is 43, obtain the position coordinates on the characteristic curve grid node after leading edge shock and shock wave ripple and flow parameter, wherein, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed and local flow direction angle, the position coordinates on characteristic curve grid node on leading edge shock 10 '-13 '-41 can represent leading edge shock profile,It is wing axis symmetric reference flow field by leading edge shock 10 '-13 '-41, tip revolving body bus 43 and straight line 41-39 area defined.

S3. given aircraft precursor costa, inlet lip molded line, air intake duct sweepforward side plate costa and the leading edge of a wing line drop shadow curve in bottom transverse cross section; From precursor costa, inlet lip molded line and air intake duct sweepforward side plate costa, hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field carries out streamlined impeller, generate hypersonic aircraft precursor-air intake duct integration configuration; From leading edge of a wing line, wing axis symmetric reference flow field carries out streamlined impeller, generate wing configuration; Hypersonic aircraft precursor-air intake duct integration configuration and wing configuration collectively constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

2. hypersonic aircraft precursor according to claim 1, air intake duct and wing rider integrated design method, it is characterized in that: in step S1.1, the selection principle of the cross section 12 at inlet lip place is this cross section with point 10 along the distance of X-direction more than aircraft precursor length.

3. hypersonic aircraft precursor according to claim 1, air intake duct and wing rider integrated design method, it is characterized in that: in step S1.1, position coordinates is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, flow parameter includes local static pressure, local density, local speed, local flow direction angle, and the position coordinates on characteristic curve grid node on leading edge shock 15 can represent leading edge shock profile. Described characteristic curve grid node is the intersection point of left lateral characteristic curve and right lateral characteristic curve.

4. hypersonic aircraft precursor according to claim 1, air intake duct and wing rider integrated design method, it is characterised in that in step S1.2, utilizes the alternative manner estimated-correct to solve the position of reflected shock wave 13-18, and its method is as follows:

The starting point of reflected shock wave 13-18 is point 13, the intersection point of reflected shock wave 13-18 and left lateral characteristic curve is called shock point, the position solving reflected shock wave 13-18 is to solve for the coordinate figure of all shock points, until the intersection point 18 of reflected shock wave 13-18 and tip revolving body bus 10-11;

For the shock point 22 and 23 that any two on reflected shock wave 13-18 is adjacent, it is defined as upstream shock point 22 near the shock point 22 putting 13, the shock point 23 of point of distance 13 is defined as downstream shock point 23, the coordinate figure of upstream shock point 22 the coordinate figure method solving downstream shock point 23 is as described below:

Characteristic curve grid node is the intersection point of left lateral characteristic curve and right lateral characteristic curve, the position coordinates of characteristic curve grid node and flow parameter all can pass through have rotation characteristic line method to solve and obtain, the position coordinates of characteristic curve grid node is the characteristic curve grid node coordinate figure on axial coordinate axle X and coordinate figure on radial coordinate axle Y under cylindrical-coordinate system, and flow parameter includes local static pressure, local density, local speed, local flow direction angle;

Shown in predicting equation in the alternative manner estimated-correct such as formula (1), shown in the iterative equation of correction such as formula (2):

$r_{i+1}^0=r_i+tan(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)\mathrm{\Δ}x---\left(1\right)$

$r_{i+1}^n=r_i+tan\[\frac{(\mathrm{\π}-({\mathrm{\β}}_i-{\mathrm{\θ}}_{i,1}\left)\right)+(\mathrm{\π}-({\mathrm{\β}}_{i+1}^{n-1}-{\mathrm{\θ}}_{i+1,1}^{n-1}\left)\right)}{2}\]\mathrm{\Δ}x---\left(2\right)$

Wherein, x is the shock point coordinate at the axial coordinate axle of cylindrical-coordinate system, and r is the shock point coordinate at the radial coordinate axle of cylindrical-coordinate system, r_iValue for the radial coordinate axle in cylindrical-coordinate system of upstream shock point 22, i is the Position Number of shock point, Δ x is downstream shock point 23 and upstream shock point 22 difference in X-direction, and β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction;It is the r value after downstream shock point 23 is estimated,It is that downstream shock point 23 corrects r value obtained after n time; θ_i,1It is the local flow direction angle θ value of the wavefront of upstream shock point 22,It is the downstream shock point 23 local flow direction angle θ value that corrects wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; β_iIt is the β value of upstream shock point 22,It is that downstream shock point 23 corrects β value obtained after n-1 time,Solved by formula (3) and obtain:

$\tan \left({\mathrm{\θ}}_{i+1,1}^{n-1}-{\mathrm{\θ}}_{i+1,2}\right)=2\cot \mathrm{\β}\frac{{\left(M_{i+1,1}^{n-1}\right)}^2{\sin }^2{\mathrm{\β}}_{i+1}^{n-1}-1}{{\left(M_{i+1,1}^{n-1}\right)}^2\left(\mathrm{\γ}+\cos \left(2{\mathrm{\β}}_{i+1}^{n-1}\right)\right)+2}---\left(3\right)$

Wherein,WithRespectively downstream shock point 23 corrects local Mach number M value and the local flow direction angle θ value of wavefront obtained after n-1 time,Obtained by the θ value linear interpolation of the point 20 on left lateral characteristic curve and point 21; θ_i+1,2It is the local flow direction angle θ value after the ripple of downstream shock point 23, θ_i+1,2It is known conditions, θ_i+1,2Obtain according to the flow direction after reflected shock wave 13-18 ripple is angular distribution.

5. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 4, it is characterized in that, in step S1.2, described oblique shock wave relational expression is utilized to solve after reflected shock wave ripple the formula of flow parameter as shown in (4)～(8):

$tan\left(\mathrm{\Δ}\mathrm{\θ}\right)=2\cot \mathrm{\β}\frac{M_1^2{\sin }^2\mathrm{\β}-1}{M_1^2(\mathrm{\γ}+cos2\mathrm{\β})+2}---\left(4\right)$

Δ θ=θ₁-θ₂(5)

$\frac{P_2}{P_1}=\frac{2\mathrm{\γ}}{\mathrm{\γ}+1}(M_1^2{\sin }^2\mathrm{\β}-\frac{\mathrm{\γ}-1}{2\mathrm{\γ}})---\left(6\right)$

$\frac{{\mathrm{\ρ}}_1}{{\mathrm{\ρ}}_2}=\frac{2}{\mathrm{\γ}+1}(\frac{1}{M_1^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{2})---\left(7\right)$

$\frac{V_2}{V_1}=\frac{sin\mathrm{\β}}{sin\[\mathrm{\β}-\mathrm{\Δ}\mathrm{\θ}\]}(\frac{2}{(\mathrm{\γ}+1)M^2{\sin }^2\mathrm{\β}}+\frac{\mathrm{\γ}-1}{\mathrm{\γ}+1})---\left(8\right)$

Wherein, β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle of shock wave and velocity of wave front direction, and Δ θ is the local flow-deviation angle of reflected shock wave, θ₁It is the local flow direction angle of reflected shock wave wavefront, M₁It is the local Mach number of reflected shock wave wavefront, P₁It is the local static pressure of reflected shock wave wavefront, ρ₁It is the local density of reflected shock wave wavefront, V₁It is the local speed of reflected shock wave wavefront, θ₂It is the local flow direction angle after reflected shock wave ripple, P₂It is the local static pressure after reflected shock wave ripple, ρ₂It is the local density after reflected shock wave ripple, V₂It is the local speed after reflected shock wave ripple.

6. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 5, it is characterized in that, in step S1.2, the angle in the reflected shock wave velocity of wave front direction at the infinitesimal of the reflected shock wave at shock point place and shock point place is the reflected shock wave local Angle of Shock Waves β in shock point position, and the angle of the reflected shock wave velocity of wave front direction at shock point place and the axial coordinate axle of cylindrical-coordinate system is the reflected shock wave wavefront flow direction angle θ in shock point position₁, after the reflected shock wave ripple at shock point place, velocity attitude is reflected shock wave flow direction angle θ after the ripple of shock point position with the angle of the axial coordinate axle of cylindrical-coordinate system₂, behind the reflected shock wave velocity of wave front direction at shock point place and the reflected shock wave ripple at shock point place, the angle of velocity attitude is the reflected shock wave local flow-deviation angle Δ θ in shock point position.

7. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 1, it is characterised in that the method for step S3 is:

Given aircraft body costa drop shadow curve 47-48-49-50-51, this curve is as the aircraft body costa drop shadow curve in bottom transverse cross section 40, wherein, point 48 and 50 is the intersection point of aircraft body costa drop shadow curve and the shock wave contour line 45 on the cross section 12 at inlet lip place, and point 47 and point 51 are the intersection points of aircraft body costa drop shadow curve and the shock wave contour line 46 on bottom transverse cross section 40;

The center of circle of shock wave contour line 45 and shock wave contour line 46 overlaps, the left side ray 44-52 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve 47-48-49-50-51 meet at a little 52, and meet at a little 54 with shock wave contour line 45, the right side ray 44-53 sent by the center of circle 44 of shock wave contour line 45 and shock wave contour line 46 and aircraft body costa drop shadow curve 47-48-49-50-51 meet at a little 53, and meet at a little 55 with shock wave contour line 45, left side ray 44-52, the angle in right side ray 44-53 and the longitudinally asymmetric face 64 of aircraft is Φ value,Curved section between point 52, point 49 and point 53, as the precursor costa drop shadow curve in bottom transverse cross section 40, is called precursor costa drop shadow curve 52-49-53; The curved section between curved section and point 53 and point 51 between point 47 and point 52, as the leading edge of a wing line drop shadow curve in bottom transverse cross section 40, is called leading edge of a wing line projection curve 47-52 and leading edge of a wing line projection curve 53-51; Curved section between point 54 and point 55, as the inlet lip molded line drop shadow curve in bottom transverse cross section 40, is called inlet lip molded line drop shadow curve 54-55;

Application free-streamline method, by precursor costa drop shadow curve 52-49-53, inlet lip molded line drop shadow curve 54-55 and leading edge of a wing line projection curve 47-52 and leading edge of a wing line projection curve 53-51, calculate precursor costa, inlet lip molded line and leading edge of a wing line respectively;

From precursor costa and inlet lip molded line, at hypersonic aircraft precursor-air intake duct integral shaft symmetric reference, flow field carries out streamlined impeller, solve all streamlines through precursor costa and inlet lip molded line, until air intake port cross section 34 position, and then obtain air intake port molded line and closed loop curve 56-57-58-59-60-61; From leading edge of a wing line, in wing axis symmetric reference flow field, carry out streamlined impeller, solve all streamlines through leading edge of a wing line, until position, bottom transverse cross section 40, and then obtain trailing edge line 47-62 and 63-51;

All streamline setting-outs through precursor costa are become stream interface 65, all streamline setting-outs through inlet lip molded line is become stream interface 66, and forms hypersonic aircraft precursor-air intake duct integration configuration plus air intake duct sweepforward side plate 67,65,66 and 67;

All streamline setting-outs on the left of leading edge of a wing line are become stream interface 68, and all streamline setting-outs on the right side of leading edge of a wing line are become stream interface 69, upper surface application free-streamline method generates, and forms wing configuration, and described wing configuration includes port wing 70 and starboard wing 71;

Described hypersonic aircraft precursor-air intake duct integration configuration and wing configuration constitute hypersonic aircraft precursor-air intake duct-wing rider integration configuration.

8. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 7, it is characterised in that in step S3, free-streamline method the implementation generating precursor costa and inlet lip molded line is as follows:

Set up an office 72 is a discrete point on precursor costa drop shadow curve 52-49-53 or inlet lip molded line drop shadow curve 54-55, intersect at a point with leading edge shock 10-13 with through point 72 straight line 73 parallel with the axial coordinate axle X of cylindrical-coordinate system, this intersection point is the point on precursor costa or inlet lip molded line, being referred to as precursor leading edge point or inlet lip point 74, straight line 74-72 is the free-stream line through precursor leading edge point or inlet lip point 74;

From precursor leading edge point or inlet lip point 74s, using the position coordinates on each characteristic curve grid node in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field and flow parameter as known conditions, streamlined impeller method is utilized to solve streamline 75, until air intake port cross section 34, the streamline 75 distal point on air intake port cross section 34 is the point on air intake port molded line, by title air intake port point 76;

Use above-mentioned same procedure, solve and obtain all precursor leading edge point and inlet lip point, and all streamlines through precursor leading edge point and inlet lip point, and obtain the air intake port point corresponding with precursor leading edge point and the air intake port point corresponding with inlet lip point; All precursor leading edge point composition precursor costa, all inlet lip points composition inlet lip molded line, the upper wall surface 56-57-58 of all air intake port point composition air intake port molded line corresponding with precursor leading edge point, the lower wall surface 61-60-59 of all air intake port point composition air intake port molded line corresponding with inlet lip point.

9. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 8, it is characterised in that in step S3, the described implementation by free-streamline method generation leading edge of a wing line and generation trailing edge line is as follows:

Set up an office 77 is a discrete point on leading edge of a wing line projection curve, intersect at a point with leading edge shock 10 '-13 '-41 with through point 77 straight line 78 parallel with the axial coordinate axle X of cylindrical-coordinate system, this intersection point is the point on leading edge of a wing line, it is called leading edge of a wing point 79, and straight line 79-77 is the free-stream line through leading edge point 79;

From leading edge of a wing point 79s, using the position coordinates on characteristic curve grid node in wing axis symmetric reference flow field and flow parameter as known conditions, streamlined impeller method is utilized to solve streamline 80, until bottom transverse cross section 40, the streamline 80 distal point on bottom transverse cross section 40 is the point on trailing edge line, and this point is called trailing edge point 81;

Use above-mentioned same procedure, solve and obtain all leading edges of a wing point, and all streamlines through leading edge of a wing point, and obtain all trailing edges point, all port wing leading edge point and starboard wing leading edge point separately constitute port wing costa and starboard wing costa, and all port wing trailing edge points and starboard wing trailing edge point separately constitute port wing trailing edge line and starboard wing trailing edge line; Port wing costa and starboard wing costa composition leading edge of a wing line, port wing trailing edge line and starboard wing trailing edge line composition trailing edge line.

10. hypersonic aircraft precursor-air intake duct-wing rider integrated design method according to claim 9, it is characterised in that in step S3, the implementation of described air intake duct sweepforward side plate costa and air intake duct sweepforward side plate is as follows:

Air intake duct sweepforward side plate includes the costa of left plate and right plate, left plate and the right plate drop shadow curve in bottom transverse cross section 40 respectively curve 52-54 and curve 53-55, and side plate costa is arranged on reflected shock wave 13-18;

Intersecting at side plate leading edge point 83 with through point 52 straight line 82 parallel with the axial coordinate axle X of cylindrical-coordinate system with reflected shock wave 13-18, side plate leading edge point 83 is the distal point on left plate costa; Side plate leading edge point 85 is intersected at reflected shock wave 13-18 with through point 54 straight line 84 parallel with the axial coordinate axle X of cylindrical-coordinate system, side plate leading edge point 85 is the starting point on left plate costa, on two dimensional surface, side plate leading edge point 85 overlaps with the point 13 on reflected shock wave;

From side plate leading edge point 83s, the position coordinates reflected shock wave in hypersonic aircraft precursor-air intake duct integral shaft symmetric reference flow field relied in district 32 and stable region 38 on characteristic curve grid node and flow parameter are as known conditions, streamlined impeller method is utilized to solve streamline 86, until air intake port cross section 34, the streamline 86 distal point on air intake port cross section 34 is the point on air intake duct side plate trailing edge line, and this distal point is called side plate trailing edge point 87;

By above-mentioned identical method, solve and obtain all side plate leading edge point, and all streamlines through side plate leading edge point, and obtain all side plate trailing edge points; All streamline setting-outs through left plate leading edge point are become stream interface, forms left plate, all left plate leading edge point are formed left plate costa, all left plate trailing edge points are formed the left side wall 56-61 of air intake port molded line; All streamline setting-outs through right plate leading edge point are become stream interface, forms right plate, all right plate leading edge point are formed right plate costa, all right plate trailing edge points are formed the right side wall 58-59 of air intake port molded line; Left plate costa and right plate costa composition air intake duct sweepforward side plate costa.