CN104210672A - Integrated design method for hypersonic-velocity wave rider fuselage and air inlet channel - Google Patents

Abstract

The invention discloses an integrated design method for a hypersonic-velocity wave rider fuselage and an air inlet channel. A wave rider is used as a whole hypersonic-velocity aircraft fuselage and is integrated with the flow line tracking outer steering air inlet channel. The integrated design method comprises the following steps: designing a pointed-end rotary body and solving a supersonic-velocity axially symmetric flow field around the zero-incidence pointed-end rotary body; designing an inner flow and outer flow integrated axially symmetric standard flow field in the supersonic-velocity axially symmetric flow field around the zero-incidence pointed-end rotary body; carrying out flow line tracking in the inner flow and outer flow integrated axially symmetric standard flow field from a front leading edge of the wave rider fuselage and an inlet molded line of the air inlet channel to generate an integrated shape. With the adoption of the technical scheme, the lift-drag ratio property of the wave rider can be sufficiently expressed.

Description

Hypersonic rider fuselage and inlet channel integrated design method

Technical field

The present invention relates to the technical field of hypersonic aircraft Design of Aerodynamic Configuration, be specifically related to a kind of hypersonic rider fuselage and inlet channel integrated design method.

Background technology

Air-breathing hypersonic vehicle refer to flight Mach number be greater than 5, with airbreathing motor or its combination engine be major impetus, can at atmospheric envelope and the aircraft across the flight of atmospheric envelope medium-long range, its application form comprises hypersonic cruise missile, hypersonicly has multiple aircraft such as people/unmanned aerial vehicle and aerospace plane etc.

Large quantity research since the sixties in 20th century absolutely proves, the integrated design of propulsion system and body is the key realizing hypersonic flight, be one of hypersonic aircraft technology gordian technique urgently to be resolved hurrily, the core of body/Propulsion Integrated is then the integration of aircraft fuselage and inlet channel.Consider from design angle, totally difference is also existed to the requirement of the two: high lift-drag ratio is mainly to the requirement of fuselage, high effective volume, and good leading edge Aerodynamic Heating barrier propterty; Then provide effective source of the gas as much as possible with the loss of minimum flowed energy for combustion chamber to the requirement of inlet channel.Good fuselage/Propulsion Integrated configurational energy meets the integration requirement of designer to hypersonic aircraft aero-propulsive performance.

Due to the difference of the two job requirement, in a very long time, people think that integration is exactly design two high performance units respectively always, carry out simple superposition mutually and trade off to them.But integrated design problem is simply really not so, the key of restriction total performance improvement is to lack a kind of rationally efficient integrated design method.

The aerodynamic configuration of hypersonic aircraft, mainly contain rotational symmetry configuration, lifting body configuration and waverider-derived three major types, wherein waverider-derived utilizes shock wave compression principle (rider principle) to achieve the pneumatic requirement of high lift-drag ratio under hypersonic flight condition, except how taking into account the problem of the protection of leading edge Aerodynamic Heating and aeroperformance, the research of this configuration is tending towards ripe.

As shown in figures 1 and 3, the conventional approach of Waverider and inlet channel integrated design is that Waverider 1 is used as hypersonic aircraft precursor, referred to as waverider forebody derived 1, inlet channel adopts two dimensional inlet, Waverider 1 is as the precompressed compression face of inlet channel, for inlet channel provides the air-flow after precompressed compression, waverider forebody derived 1 produces leading edge shock 5, leading edge shock 5 is incident on inlet lip 2, and produce reflected shock wave 6, air-flow enters inlet channel distance piece 4, for combustion chamber provides source of the gas through the compression of leading edge shock 5, reflected shock wave 6 and inlet channel outer cover 3.In this conventional Waverider-inlet channel integrated design method, only Waverider is used as hypersonic aircraft precursor, the high lift-drag ratio characteristic of Waverider can not be given full play to.

Summary of the invention

The technical problem to be solved in the present invention is, provides a kind of hypersonic rider fuselage and inlet channel integrated design method, can give full play to the high lift-drag ratio characteristic of Waverider.

For solving the problem, the present invention adopts following technical scheme:

A kind of hypersonic rider fuselage and inlet channel integrated design method, adopt using Waverider as whole hypersonic aircraft fuselage, while with to turn to inlet channel to carry out outside streamlined impeller integrated, comprise the following steps:

Step S1, design tip gyro-rotor, and solve the supersonic speed axisymmetric flow field around zero-incidence tip gyro-rotor;

Step S2, described in the supersonic speed axisymmetric flow field of zero-incidence tip gyro-rotor, the inside and outside flow integrated rotational symmetry benchmark flow field of design;

Step S3, from rider fuselage costa and inlet mouth molded line, in described inside and outside flow integrated rotational symmetry benchmark flow field, carry out streamlined impeller, generate integrated configuration.

The present invention adopts using Waverider as whole hypersonic aircraft fuselage, referred to as rider fuselage, simultaneously with to turn to inlet channel to carry out outside streamlined impeller integrated, makes aircraft under design point, gives full play to the high lift-drag ratio characteristic of Waverider.The present invention designs a kind of axisymmetric flow field simultaneously with interior flowing and outer flowing, can be used as the benchmark flow field of rider fuselage configuration integrated with turning to inlet channel outside streamlined impeller.The present invention names this axisymmetric flow field for " inside and outside flow integrated rotational symmetry benchmark flow field ", and it comprises the inside and outside inlet channel reference flow place in flow integrated rotational symmetry benchmark flow field and the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field.By making rider fuselage costa and inlet mouth molded line share a segment type line, realize rider fuselage and being fused into one of inlet channel configuration.From rider fuselage costa, streamlined impeller is carried out in the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, and from inlet mouth molded line, carrying out streamlined impeller in the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, generating rider fuselage configuration integrated with turning to inlet channel outside streamlined impeller.Under design point, whole hypersonic aircraft fuselage has rider characteristic, and waverider forebody derived is as the precompressed compression face of inlet channel, and for inlet channel provides the air-flow after precompressed compression, rider fuselage provides high lift-drag ratio for aircraft.

Accompanying drawing explanation

Fig. 1 is the schematic three dimensional views of conventional Waverider-inlet channel integrated design scheme;

Fig. 2 is the schematic diagram of longitudinal plane of symmetry of conventional Waverider-inlet channel integrated design scheme;

Fig. 3 is the supersonic speed axisymmetric flow field schematic diagram around zero-incidence tip gyro-rotor;

Fig. 4 is that the leading edge shock in rider fuselage reference flow place and inlet channel benchmark flow field relies on district;

Fig. 5 is main compression zone and the reflected shock wave of isentropic compression between the shock wave of inlet channel benchmark flow field;

Fig. 6 is the schematic diagram solving reflected shock wave position;

Fig. 7 is the flow direction angle θ of the local Angle of Shock Waves β of reflected shock wave, reflected shock wave wavefront ₁, the flow direction angle θ after reflected shock wave ripple ₂and the definition of the local flow-deviation angle Δ θ of reflected shock wave;

Fig. 8 is that the reflected shock wave in inlet channel benchmark flow field relies on district;

Fig. 9 is the stable region in inlet channel benchmark flow field and inside and outside flow integrated rotational symmetry benchmark flow field;

Figure 10 shows the projection of shock wave in gyro-rotor bottom transverse cross section of rider fuselage costa, inlet mouth molded line and cross-sectional plane 1 and cross-sectional plane 2 position;

Figure 11 is each molded line schematic diagram of rider fuselage configuration integrated with turning to inlet channel outside streamlined impeller.

Figure 12 is the stream interface of structure rider fuselage, upper surface and bottom surface;

Figure 13 is the stream interface turning to inlet channel outside structure streamlined impeller;

Figure 14 is by the rider fuselage turning to inlet channel to form outside rider fuselage and streamlined impeller configuration integrated with turning to inlet channel outside streamlined impeller;

Figure 15 is the design diagram of rider fuselage costa, streamline and trailing edge line;

Figure 16 is the design diagram of inlet mouth molded line, streamline and air intake port molded line.

In figure, 1 represents waverider forebody derived; 2 represent normal inlet lip; 3 represent inlet channel outer cover; 4 represent inlet channel 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; 8 represent tip gyro-rotor summit; X represents the axial coordinate axle of cylindrical-coordinate system; Y represents the radial coordinates axle of cylindrical-coordinate system; 11 represent tip gyro-rotor bus; The end wall millet cake of 12 expression tip gyro-rotor buses 11; 13 represent tip gyro-rotor bottom transverse cross section; 14 represent the leading edge shock around tip gyro-rotor; 15 represent the distal point of tip gyro-rotor leading edge shock 14 at cross-sectional plane 13 place; 16 represent the unique point chosen on tip gyro-rotor bus; 17 represent through the some left lateral characteristic curve of 16 and the intersection point of leading edge shock; 18 represent the streamline through point 17; 19 represent the distal point of streamline 18 at cross-sectional plane 13 place; 20 represent the rider fuselage reference flow place surrounded by 17-15-19; 21 represent that the leading edge shock in the inlet channel benchmark flow field surrounded by 8-17-16 relies on district; 22 represent by passing through the some right lateral characteristic curve of 17 and the wall intersection point of gyro-rotor bus; The wall intersection point of 23 expression reflected shock waves and gyro-rotor bus; The 24 main compression zones representing isentropic compression between the inlet channel benchmark flow field shock wave that surrounded by 16-17-23; 27 and 28 represent the adjacent feature wire grid node on same left lateral characteristic curve; 29 and 30 represent the adjacent shock point on reflected shock wave 17-23; The reflected shock wave at 31 expression point 30 places; The reflected shock wave velocity of wave front direction at 32 expression point 30 places; Velocity reversal after the reflected shock wave ripple at 33 expression point 30 places; The local Angle of Shock Waves β of 34 expression point 30 place's reflected shock waves; 35 represent point 30 place reflected shock wave wavefront air flow line angle θ ₁; Air flow line angle θ after 36 expression point 30 place's reflected shock wave ripples ₂; 37 represent point 30 place's reflected shock wave local flow-deviation angle Δ θ; The 38 wall intersection points representing right lateral characteristic curve through point 23 and inlet channel reference flow wall outside the venue; 39 represent that the reflected shock wave in the inlet channel benchmark flow field surrounded by a 17-38-23 relies on district; 40 represent the inlet channel benchmark flow field internal face curve on the right side of point 23; 41 represent the inlet channel reference flow wall curve outside the venue on the right side of point 38; 42 distal points representing inlet channel reference flow wall outside the venue; 43 represent through the some right lateral characteristic curve of 42 and the wall intersection point of inlet channel benchmark flow field internal face; 44 stable regions representing the inlet channel benchmark flow field surrounded by 23-38-42-43; 45 represent the cross-section location through point 17; 46 represent air intake port cross-sectional plane, are the cross-sectional plane through putting 42; 47 represent the outline line of leading edge shock 14 in cross-sectional plane 45 position; 48 represent the outline line of leading edge shock 14 in cross-sectional plane 13 position; The left end point of 49 expression rider fuselage costas; The left end point of 20 expression inlet mouth molded line; 51 represent the mid point of rider fuselage costa and the mid point of inlet mouth molded line; The right endpoint of 52 expression inlet mouth molded line; The right endpoint of 53 expression rider fuselage costas; The lower extreme point of 54 expression inlet mouth molded line; 55 represent the drop shadow curve of rider fuselage costa 48-49-50-51-52 at cross-sectional plane 13; 56 represent inlet mouth molded line 49-50-51-53, and 57 represent the outline line of tip gyro-rotor at cross-sectional plane 13; 58 represent rider fuselage trailing edge line; 59 represent air intake port molded line; The stream interface of 60 expression rider fuselages, is the lower surface of rider fuselage; 61 represent the curve 48-49-53-51-52 that is made up of the segment of curve 51-52 tri-sections of the segment of curve 48-49 of rider fuselage costa, the segment of curve 49-53-51 of inlet mouth molded line and the rider fuselage costa drop shadow curve at cross-sectional plane 13; The upper surface of 62 expression rider fuselages; The bottom surface of 63 expression rider fuselages; 64 represent the point of rider fuselage costa in the drop shadow curve of cross-sectional plane 13; 65 represent through point 64 and the straight line parallel with the axial coordinate axle of cylindrical-coordinate system; 66 represent the point on rider fuselage costa, are the intersection point of straight line 65 and leading edge shock 14; 67 represent the streamline through point 66 in the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field; 68 represent the distal point of streamline 67 at cross-sectional plane 13; 69 represent the point of inlet mouth molded line in the drop shadow curve of cross-sectional plane 13; 70 represent through point 69 and the straight line parallel with the axial coordinate axle of cylindrical-coordinate system; 71 represent the point on inlet mouth molded line, are the intersection point of straight line 70 and leading edge shock 14; 72 represent the streamline through point 71 in the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field; 73 represent the distal point of streamline 72 on air intake port cross-sectional plane 68.

Detailed description of the invention

The invention provides a kind of hypersonic rider fuselage and inlet channel integrated design method, adopt using Waverider as whole hypersonic aircraft fuselage, referred to as rider fuselage, simultaneously with to turn to inlet channel to carry out outside streamlined impeller integrated, comprise the following steps:

Step S1, design tip gyro-rotor, and solve the supersonic speed axisymmetric flow field around zero-incidence tip gyro-rotor.

As shown in Figure 3, if curve 11 is buses of tip gyro-rotor, the starting point of bus is point 8, the distal point of bus is point 12, this tip gyro-rotor is under the effect of zero-incidence and Supersonic Stream 7, can produce leading edge shock 14, the distal point of leading edge shock 14 on tip gyro-rotor bottom transverse cross section 13 is point 15.

Using the condition of Supersonic Stream 7 as input parameter, inlet flow conditions comprises incoming flow Mach, incoming flow static pressure, incoming flow static temperature, utilization has revolves characteristic line method, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", solve the supersonic speed axisymmetric flow field around zero-incidence tip gyro-rotor, obtain the position coordinate on the characteristic curve grid node after leading edge shock and shock wave ripple and flow parameter, position coordinate is the coordinate figure of characteristic curve grid node under cylindrical-coordinate system on axial coordinate axle X and the coordinate figure on radial coordinates axle Y, under cylindrical-coordinate system, axial coordinate axle X is the axis of tip gyro-rotor, flow parameter comprises local static pressure, local density, local speed, local flow direction angle, position coordinate on characteristic curve grid node on leading edge shock 14 can indicate leading edge shock profile.

Step S2, described in the supersonic speed axisymmetric flow field of zero-incidence tip gyro-rotor, the inside and outside flow integrated rotational symmetry benchmark flow field of design.

The concrete steps designing this benchmark flow field comprise the steps:

S2.1, as shown in Figure 4, tip gyro-rotor bus 8-12 gets a little 16, and the selection principle of point 16 is left sides that the intersection point of left lateral characteristic curve and the leading edge shock 14 sent by point 16 must be positioned at leading edge shock distal point 15, otherwise cannot design reflectivity shock wave.

By the position coordinate of Supersonic Stream condition and wall 8-16, inlet flow conditions comprises free stream Mach number, incoming flow static pressure, incoming flow static temperature, utilization has revolves characteristic line method, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", solve through putting the left lateral characteristic curve of 16 and the intersection point 17 of leading edge shock, and the leading edge shock solving the inlet channel benchmark flow field surrounded by 8-17-16 relies on district 21, supersonic speed axisymmetric flow field around zero-incidence tip gyro-rotor is divided into two regions by the streamline 18 passing through point 17, i.e. region outside streamline 18 and the region inside streamline 18, outside streamline 18, the region 20 surrounded by a 17-15-19 is as rider fuselage reference flow place, region inside streamline 18 is used for the design in inlet channel benchmark flow field, point 19 is distal points of streamline 18.

S2.2, as shown in Figure 5, utilization has revolves characteristic line method, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics ", the left Crow of M.J.; J.D. Huffman; National Defense Industry Press, p138-195 in 1984 ", solve through putting the right lateral characteristic curve of 17 and the wall intersection point 22 of gyro-rotor bus, and solve by the flow field of 16-17-22 institute enclosing region; Point 17 is as the initial point of the reflected shock wave in inlet channel benchmark flow field, flow direction angle distribution after given reflected shock wave 17-23 ripple, utilize the alternative manner estimating-correct, solve the position of reflected shock wave 17-23, then utilize oblique shock wave relational expression to solve flow parameter after ripple.By a 16-17-23 surround the main compression zone of region 24 as isentropic compression between the shock wave of inlet channel benchmark flow field.Wherein, flow direction angle is the angle of the axial coordinate axle of flow direction and cylindrical-coordinate system.

The position concrete grammar that the alternative manner that described utilization is estimated-corrected solves reflected shock wave 17-23 is as follows.

As shown in Figure 6, fine line in Fig. 6 represents left lateral characteristic curve, represented by dotted arrows right lateral characteristic curve, hollow node on behalf characteristic curve grid node, the initial point of reflected shock wave 17-23 is point 17, the intersection point of reflected shock wave 17-23 and left lateral characteristic curve referred to as shock point, such as upstream shock point 29, downstream shock point 30, the described position solving reflected shock wave 17-23 is the coordinate figure solving all shock points, until the wall intersection point 23 of reflected shock wave and gyro-rotor bus.

The coordinate figure method being solved downstream shock point 30 by the coordinate figure of upstream shock point 29 is as described below.

In Fig. 6, the position coordinate of characteristic curve grid node and flow parameter all can revolve characteristic line method and solve and obtain by having, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics "; the left Crow of M.J.; J.D. Huffman; National Defense Industry Press; 1984; p138-195 ", position coordinate is the coordinate figure of characteristic curve grid node under cylindrical-coordinate system on axial coordinate axle X and the coordinate figure on radial coordinates axle Y, and flow parameter comprises local static pressure, local density, local speed, local flow direction angle.

Described predicting equation is such as formula shown in (1), and the iterative equation of correction is such as formula shown in (2).

$r_{i+1}^0=r_i+\tan (\mathrm{\π}-{\mathrm{\β}}_i)\mathrm{\Δx}---\left(1\right)$

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

Wherein, x is the coordinate of the axial coordinate axle of cylindrical-coordinate system, and r is the coordinate of the radial coordinates axle of cylindrical-coordinate system, r _ifor the value of the radial coordinates axle in cylindrical-coordinate system of upstream shock point 29, i is the Position Number of shock point, and 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, described local Angle of Shock Waves is the angle in shock wave and velocity of wave front direction the r value after downstream shock point 30 is estimated, the r value that downstream shock point 30 corrects that after n time, gained arrives, β _ithe β value of upstream shock point 29, the β value that downstream shock point 30 corrects that after n-1 time, gained arrives, solved by formula (3) and obtain.

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

Wherein, M is the Mach number of reflected shock wave in locality, the local flow direction angle θ value that downstream shock point 30 corrects the wavefront that gained arrives after n-1 time, obtained by the θ value linear interpolation of the point 27 on left lateral characteristic curve and point 28; θ _i+1,1the local flow direction angle θ value after the ripple of downstream shock point 30, θ _i+1,1be known conditions, it can obtain according to the flow direction angle distribution after reflected shock wave 17-23 ripple.

The described formula utilizing oblique shock wave relational expression to solve flow parameter after reflected shock wave ripple is as shown in (4) ~ (8).

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

Δθ＝θ ₁-θ2 (5)

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

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

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

Wherein, β is the local Angle of Shock Waves of reflected shock wave, and described local Angle of Shock Waves is the angle in shock wave and velocity of wave front direction, and Δ θ is the local flow-deviation angle of reflected shock wave, θ ₁the local flow direction angle of reflected shock wave wavefront, M ₁the local Mach number of reflected shock wave wavefront, P ₁the local static pressure of reflected shock wave wavefront, ρ ₁the local density of reflected shock wave wavefront, V ₁the local speed of reflected shock wave wavefront, θ ₂the local flow direction angle after reflected shock wave ripple, P ₂the local static pressure after reflected shock wave ripple, ρ ₂the local density after reflected shock wave ripple, V ₂the 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 wavefront ₁, the flow direction angle θ after reflected shock wave ripple ₂and the definition of the local flow-deviation angle Δ θ of reflected shock wave as shown in Figure 7, reflected shock wave 31 in shock point 30 position is reflected shock wave local Angle of Shock Waves β in shock point 30 position with the angle 34 in reflected shock wave velocity of wave front direction 32, and the velocity reversal 32 of reflected shock wave wavefront is the wavefront flow direction angle θ of reflected shock wave in shock point 30 position with the angle 35 of the axial coordinate axle of cylindrical-coordinate system ₁, the velocity reversal 33 after reflected shock wave ripple is reflected shock wave flow direction angle θ after the ripple of shock point 30 position with the angle 36 of the axial coordinate axle of cylindrical-coordinate system ₂, the angle 37 of the velocity reversal 33 after the velocity reversal 32 of reflected shock wave wavefront and reflected shock wave ripple is reflected shock wave local flow-deviation angle Δ θ in shock point 30 position.

S2.3, as shown in Figure 8, utilization has revolves characteristic line method, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics "; the left Crow of M.J.; J.D. Huffman; National Defense Industry Press; 1984; p138-195 ", by the flow parameter after reflected shock wave 17-23 ripple, solve the wall intersection point 38 of right lateral characteristic curve through putting 23 and inlet channel reference flow wall outside the venue, thus obtain inlet channel reference flow wall 17-38 section outside the venue, and the reflected shock wave solving the inlet channel benchmark flow field surrounded by a 17-38-23 relies on district 39.

S2.4, as shown in Figure 9, first the inlet channel benchmark flow field internal face curve 40 on the right side of set point 23 and the Mach Number Distribution on this curve, make the angle of contingence of curve 40 in point 23 position overlap with local flow direction angle, to ensure that shock wave 17-23 is in point 23 position no reflection simultaneously, utilization has revolves characteristic line method, have and revolve the known technology that characteristic line method is this area, specifically can see " " gas kinetics ", M.J. left Crow, J.D. Huffman, National Defense Industry Press, 1984, p138-195 ", by the Mach Number Distribution on wall curve 40 and this curve, inlet channel reference flow on the right side of solution point 38 wall curve 41 outside the venue, solve the stable region 44 in the inlet channel benchmark flow field surrounded by a 23-38-42-43, point 42 is distal points of outside wall surface curve 41, and the wall intersection point 43 of the right lateral characteristic curve solved through putting 42 and inlet channel benchmark flow field internal face.Wherein, cut angle the angle of the axial coordinate axle being curve near tangent and cylindrical-coordinate system.

Rotational symmetry benchmark flow field is as shown in Figure 9 a kind of axisymmetric flow field simultaneously with interior flowing and outer flowing, and the present invention names this axisymmetric flow field for " inside and outside flow integrated rotational symmetry benchmark flow field ".Inside and outside flow integrated rotational symmetry benchmark flow field is made up of five flow field regions, the stable region 44 that namely leading edge shock in rider fuselage reference flow place 20, inlet channel benchmark flow field relies on the main compression zone 24 of isentropic compression between district 21, inlet channel benchmark flow field shock wave, the reflected shock wave in inlet channel benchmark flow field relies on district 39 and inlet channel benchmark flow field.Rider fuselage reference flow place 20 constitutes the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, and the region, 44 4, stable region that the leading edge shock in inlet channel benchmark flow field relies on the main compression zone 24 of isentropic compression between district 21, inlet channel benchmark flow field shock wave, the reflected shock wave in inlet channel benchmark flow field relies on district 39 and inlet channel benchmark flow field constitutes the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field.

Step S3, from rider fuselage costa and inlet mouth molded line, in described inside and outside flow integrated rotational symmetry benchmark flow field, carry out streamlined impeller, generate integrated configuration.

As shown in Figure 10,11, given open loop curve 55, curve 55 is as the drop shadow curve of rider fuselage costa 49-50-51-52-53 at cross-sectional plane 13, referred to as rider fuselage costa drop shadow curve 55, two, the left and right end points of rider fuselage costa is the intersection point 49 and 53 of the shock wave 48 of it and cross-sectional plane 13 position respectively; Given closed loop curve 56, curve 56 is as the drop shadow curve of inlet mouth molded line 50-51-52-54 at cross-sectional plane 13, referred to as inlet mouth molded line drop shadow curve 56, two, the left and right end points of inlet mouth molded line is the intersection point 50 and 52 of the shock wave 47 of it and cross-sectional plane 45 position respectively, and two end points up and down of inlet mouth molded line are point 51 and 54 respectively.

Rider fuselage costa and inlet mouth molded line link together by sharing one section of curve 50-51-52.

By rider fuselage costa drop shadow curve 55, calculate rider fuselage costa 49-50-51-52-53, by inlet mouth molded line drop shadow curve 56, calculate inlet mouth molded line 50-51-52-54.

From rider fuselage costa, carry out streamlined impeller in the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, solve all streamlines through rider fuselage costa, until cross-sectional plane 13 position, and then obtain rider fuselage trailing edge line 58.From inlet mouth molded line, streamlined impeller is carried out in the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, solve all streamlines through inlet mouth molded line, until air intake port cross-sectional plane 46 position, and then obtain air intake port molded line 59.

As shown in Figure 11,12, all streamline setting-outs through rider fuselage costa are become stream interface, using the lower surface 60 of stream interface as rider fuselage, using the curve 49-50-54-52-53 be made up of segment of curve 49-50, the segment of curve 50-54-52 of inlet mouth molded line of rider fuselage costa, the segment of curve 52-53 tri-sections of rider fuselage costa and its plane formed in the drop shadow curve 61 of cross-sectional plane 13 upper surface 62 as rider fuselage, using the plane that is made up of drop shadow curve 61 and rider fuselage trailing edge line 58 bottom surface 63 as rider fuselage; Upper surface 62, lower surface 60 and bottom surface 63 constitute rider fuselage.

As shown in Figure 11,13, all streamline setting-outs through inlet mouth molded line are become stream interface, and stream interface constitutes outside streamlined impeller and turns to inlet channel.

As shown in figure 14, inlet channel is turned to constitute rider fuselage configuration integrated with turning to inlet channel outside streamlined impeller outside described rider fuselage and described streamlined impeller.

The implementation of described rider fuselage costa and trailing edge line is as described below.

As shown in figure 15, by the coordinate figure of the point 64 in rider fuselage costa drop shadow curve, revolve characteristic line method and solve according to having the leading edge shock shape around zero-incidence tip gyro-rotor obtained, with with the parallel straight line of the axial coordinate axle of cylindrical-coordinate system 65 and leading edge shock crossing through point 64, intersection point 66 is the points on rider fuselage costa, is called for short rider fuselage leading edge point 66, from rider fuselage leading edge point 66s, using the position coordinate on characteristic curve grid node in the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field and flow parameter as known conditions, position coordinate is the coordinate figure of characteristic curve grid node under cylindrical-coordinate system on axial coordinate axle X and the coordinate figure on radial coordinates axle Y, flow parameter comprises local static pressure, local density, local speed, local flow direction angle, utilize streamlined impeller method, streamlined impeller method is the known technology of this area, specifically can see " " 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 67, until cross-sectional plane 13, streamline 67 is the points on rider fuselage trailing edge line at the distal point 68 in bottom transverse cross section 13, be called for short rider fuselage trailing edge point 68, by same procedure, solve and obtain all rider fuselage leading edge point, and through all streamlines of rider fuselage leading edge point, and obtain all rider fuselage trailing edge points, all rider fuselage leading edge point composition rider fuselage costas, all rider fuselage trailing edge point composition rider fuselage trailing edge lines.

The implementation of described inlet mouth molded line and air intake port molded line is as described below.

As shown in figure 16, by the coordinate figure of the point 69 in inlet mouth molded line drop shadow curve, revolve characteristic line method and solve according to having the leading edge shock shape around zero-incidence tip gyro-rotor obtained, with with the parallel straight line of the axial coordinate axle of cylindrical-coordinate system 70 and leading edge shock crossing through point 69, intersection point 71 is the points on inlet mouth molded line, is called for short inlet mouth point 71, from inlet mouth point 71s, using the position coordinate on each characteristic curve grid node in the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field and flow parameter as known conditions, position coordinate is the coordinate figure of characteristic curve grid node under cylindrical-coordinate system on axial coordinate axle X and the coordinate figure on radial coordinates axle Y, flow parameter comprises local static pressure, local density, local speed, local flow direction angle, utilize streamlined impeller method, utilize streamlined impeller method, streamlined impeller method is the known technology of this area, specifically can see " " 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 72, until air intake port cross-sectional plane 46, streamline 72 is the points on air intake port molded line at the distal point 73 of air intake port cross-sectional plane 46, be called for short air intake port point 73, by same procedure, solve and obtain all inlet mouths point, and through all streamlines of inlet mouth point, and obtain all air intake ports point, all inlet mouth point composition inlet mouth molded line, all air intake port point composition air intake port molded line.

The detailed description of the invention of just this invention described in specification sheets.Although describe embodiments of the present invention by reference to the accompanying drawings, in this area, those skilled in the art can make various distortion or amendment within the scope of the appended claims.

Claims (3)

1. hypersonic rider fuselage and an inlet channel integrated design method, is characterized in that, adopts using Waverider as whole hypersonic aircraft fuselage, simultaneously with to turn to inlet channel to carry out outside streamlined impeller integrated, comprises the following steps:

Step S1, design tip gyro-rotor, and solve the supersonic speed axisymmetric flow field around zero-incidence tip gyro-rotor;

Step S2, described in the supersonic speed axisymmetric flow field of zero-incidence tip gyro-rotor, the inside and outside flow integrated rotational symmetry benchmark flow field of design;

Step S3, from rider fuselage costa and inlet mouth molded line, in described inside and outside flow integrated rotational symmetry benchmark flow field, carry out streamlined impeller, generate integrated configuration.

2. a kind of hypersonic rider fuselage as claimed in claim 1 and inlet channel integrated design method, it is characterized in that, inside and outside flow integrated rotational symmetry benchmark flow field comprises: rider fuselage reference flow place (20), the leading edge shock in inlet channel benchmark flow field relies on district (21), the main compression zone (24) of isentropic compression between the shock wave of inlet channel benchmark flow field, the reflected shock wave in inlet channel benchmark flow field relies on the stable region (44) in district (39) and inlet channel benchmark flow field, wherein, rider fuselage reference flow place (20) constitutes the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, the leading edge shock in inlet channel benchmark flow field relies on district (21), the main compression zone (24) of isentropic compression between the shock wave of inlet channel benchmark flow field, reflected shock wave dependence district (39) in inlet channel benchmark flow field and the region, (44) four, stable region in inlet channel benchmark flow field constitute the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field.

3. a kind of hypersonic rider fuselage as claimed in claim 2 and inlet channel integrated design method, it is characterized in that, step S3 is specially:

From rider fuselage costa, streamlined impeller is carried out in the rider fuselage reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, solve all streamlines through rider fuselage costa, until tip gyro-rotor bottom transverse cross section (13) position, and then obtain rider fuselage trailing edge line (58);

From inlet mouth molded line, streamlined impeller is carried out in the inlet channel reference flow place in inside and outside flow integrated rotational symmetry benchmark flow field, solve all streamlines through inlet mouth molded line, until air intake port cross-sectional plane (46) position, and then obtain air intake port molded line (59);

Rider fuselage costa and inlet mouth molded line link together by sharing one section of curve (50-51-52).