CN101813027A - Bump air inlet method for realizing integration of unequal-strength wave system with forebody - Google Patents
Bump air inlet method for realizing integration of unequal-strength wave system with forebody Download PDFInfo
- Publication number
- CN101813027A CN101813027A CN 201010134882 CN201010134882A CN101813027A CN 101813027 A CN101813027 A CN 101813027A CN 201010134882 CN201010134882 CN 201010134882 CN 201010134882 A CN201010134882 A CN 201010134882A CN 101813027 A CN101813027 A CN 101813027A
- Authority
- CN
- China
- Prior art keywords
- shock wave
- angle
- intake duct
- bulge
- wave
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 19
- 230000010354 integration Effects 0.000 title abstract 2
- 230000035939 shock Effects 0.000 claims abstract description 79
- 238000013461 design Methods 0.000 claims abstract description 61
- 238000011084 recovery Methods 0.000 claims abstract description 25
- 238000007906 compression Methods 0.000 claims abstract description 8
- 230000008676 import Effects 0.000 claims description 38
- 238000002604 ultrasonography Methods 0.000 claims description 7
- 238000004458 analytical method Methods 0.000 claims description 3
- 238000013459 approach Methods 0.000 claims description 3
- 230000015572 biosynthetic process Effects 0.000 claims 2
- 102100037651 AP-2 complex subunit sigma Human genes 0.000 abstract 2
- 101000806914 Homo sapiens AP-2 complex subunit sigma Proteins 0.000 abstract 2
- 230000006835 compression Effects 0.000 description 4
- 238000010586 diagram Methods 0.000 description 4
- 230000000694 effects Effects 0.000 description 4
- 230000014509 gene expression Effects 0.000 description 3
- 208000036366 Sensation of pressure Diseases 0.000 description 2
- 238000012938 design process Methods 0.000 description 2
- 238000009826 distribution Methods 0.000 description 2
- 239000012530 fluid Substances 0.000 description 2
- 238000000465 moulding Methods 0.000 description 2
- 238000011160 research Methods 0.000 description 2
- 238000010205 computational analysis Methods 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 238000009499 grossing Methods 0.000 description 1
- CNKHSLKYRMDDNQ-UHFFFAOYSA-N halofenozide Chemical compound C=1C=CC=CC=1C(=O)N(C(C)(C)C)NC(=O)C1=CC=C(Cl)C=C1 CNKHSLKYRMDDNQ-UHFFFAOYSA-N 0.000 description 1
- 238000005304 joining Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 239000007787 solid Substances 0.000 description 1
Images
Landscapes
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention provides a bump air inlet method for realizing integration of an unequal-strength wave system with a forebody, relating to the technical field of ultrasonic air inlet. The method comprises the following steps that: 1, an air inlet wave system adopts an external-compression double wave system structure based on an unequal-strength system, and the total pressure recovery coefficient sigma=sigma1*sigma2, wherein the sigma1 and the sigma2 are total pressure recovery coefficients of a conical shock wave and a normal shock wave; 2, a cone with a semi-cone angle delta c formed by the conical shock wave generates a second conical shock wave with second semi-cone angle beta in ultrasonic flow; 3, the bump height of the air inlet is h, the h/delta is 2-2.5 when the local boundary layer thickness is delta, the bump compressive plane generated in the step 2 is scaled to meet the requirement of an actual size; and 4, the lip of the air inlet adopts a conformal and sweepback lip design. The method realizes reduction of throat Mach number of the air inlet, improves the air inlet performance to ensure that the lip sweepback angle of the air inlet, the maximum turning angle of the bump compressive plane and the inlet normal shock wave angle are consistent, can increase the curve of the total pressure recovery coefficient, and reduces resistance coefficient.
Description
Technical field
What the present invention relates to is a kind of based on not waiting designing with the incorporate Bump intake duct of forebody of high-amplitude wave system, belongs to supersonic speed air inlet road technical field.
Background technique
The design of intake duct is one of key of fighter design.Intake duct not only will also will be considered the constraint of total arrangement and the requirement of integrated design for motor provides enough high-quality air qualities under all states when design, in addition, also must satisfy the overall stealthy requirement of fighter.For the supersonic speed air inlet road, need by a series of shock wave, the ultrasound velocity incoming flow is slowed down is subsonic flow, continues deceleration diffusion, flow direction engine again in the diffuser passage.General slant compression plate or the compression cone of adopting of traditional supersonic speed air inlet road design forms multishock, and every road and dividing plate the intake duct import is lifted away from fuselage surface by boundary layer, enters intake duct to avoid the low energy air-flow in the fuselage surface boundary layer.
No boundary layer also claims the Bump intake duct every supersonic speed air inlet road, road, be by Lockheed Martin Corporation design and on the F-35 aircraft a kind of novel intake duct of successful Application.The suction port of this intake duct is not provided with conventional fixed boundary layer every the road, but by computer design the projection piece (or bulge) of a three-dimension curved surface.The compression to air-flow is played in this bulge, and produces a pressure distribution that air stream on accompany surface is pushed away intake duct.Whole gas handling system does not have movable member, does not have the boundary layer isolating plate, does not have blow-off system and bypass system yet, has reduced by 300 pounds structure weight, has therefore reduced production and cost of use yet.
Because external consistent pneumatic design technology to the advanced person holds in close confidence, about the data of F-35 aircraft has only general report and disclosed aircraft picture, the external open source literature of relevant Bump intake duct design and performance study does not almost have.In recent years, domestic have how tame unit to carry out research work to the Bump intake duct, wherein the Bump intake duct of Chengdu Aircraft Design ﹠ Research Institute's design and having used on homemade FC-1 " brave dragon " aircraft.
Yet, the focus of Bump intake duct mainly being concentrated on the design of three-dimensional bulge compressing surface both at home and abroad, the document of having delivered is not all paid close attention to the integrated design of Bump intake duct and aircraft forebody, does not provide the selection principle of relevant integrated design parameter.
Summary of the invention
The object of the invention provides a kind of inlet throat Mach number that reduces, improve inlet characteristic, make the maximum deflection angle of inlet lip sweepback angle and bulge compressing surface, import normal shock wave angle consistent, can increase the total pressure recovery coefficient curve, reduce resistance coefficient based on not waiting high-amplitude wave system and the incorporate Bump intake duct of forebody.
The present invention adopts following technological scheme for achieving the above object:
A kind of method that realizes not waiting high-amplitude wave system and the integrated Bump intake duct of forebody, the ultrasound velocity incoming flow produces conical shock wave one at the head of bulge compressing surface, forms one normal shock wave before inlet lip;
The first step: the intake duct wave system adopts based on the external-compression type two wave system structures that do not wait high-amplitude wave system, and the total pressure recovery coefficient of intake duct multishock is σ
s=σ
1σ
2, σ wherein
1, σ
2Be respectively the total pressure recovery coefficient of conical shock wave, normal shock wave, by waiting ripple to join the ripple theory analysis by force, total pressure recovery coefficient was the highest when the ripple of twice ripple equated by force, was best wave system;
Second step: semi-cone angle is δ
cCircular cone in supersonic flow, produce the conical shock wave that semi-cone angle is β, the circle radius of conical shock wave is R, with the plane truncated cone shape shock wave of distance cone axis linear distance d, d＜R wherein, the streamline that every bit sends backward on plane taken and the conical shock wave intersection constitutes the bulge compressing surface;
The 3rd step: make that intake duct import bump height is h, local boundary layer thickness is δ, satisfy relation h/ δ=2～2.5 between intake duct import bump height h and the local boundary layer thickness δ, the bulge compressing surface that second step was generated carries out convergent-divergent, satisfies the actual size requirement;
The 4th step: inlet lip adopts conformal and the design of sweepback lip, the import antelabium is most of fits with the circular cone shock surface, the lip sweepback angle is consistent with the maximum deflection angle of bulge compressing surface, import normal shock wave angle respectively, to increase the total pressure recovery coefficient curve, to reduce resistance coefficient.
Inflow Mach number behind the normal shock wave of the present invention is not more than 0.75.
The present invention goes to intercept the conical flow flow field with plane or the curved surface that distance generates body circular cone axis different heights h, and face is terminal identical with conical tip line and axis angle as long as institute dams, and the rider build face that is then generated is similar.
The present invention calculates circular cone shock wave angle β according to the import wave system of intake duct
AwlWith circular cone semi-cone angle δ
c, determine the profile deflection angle theta then, generate the bulge compressing surface according to rider build face similar Design principle at last.
Make that the bulge width is W, bulge width W and meet following Changing Pattern apart from ratio W/d between the d and bulge compressing surface deflection angle theta, promptly when W/d 〉=10, the profile deflection angle theta approaches circular cone semi-cone angle δ
c, δ
c-θ＜1 °.
The present invention adopts technique scheme, compared with prior art has following advantage:
1, utilizes that of the present invention not wait high-amplitude wave be design method, can reduce the inlet throat Mach number, improve inlet characteristic.
2, utilize the rider body bulge compressing surface similar Design method based on the profile drift angle of the present invention, can design processes simplified, no longer need parameter d is compared and optimizes.
3, utilize bulge design height that the present invention sets up and the relation between the local boundary layer thickness, bulge and aircraft forebody can be carried out integrated design.
4, utilize conformal lip design of the present invention, can make the most of and shock surface applying of import antelabium, avoid the top overflow of lip cover.
5, utilize sweepback lip of the present invention design, make the maximum deflection angle of inlet lip sweepback angle and bulge compressing surface, import normal shock wave angle consistent, can increase total pressure recovery coefficient curve, reduction resistance coefficient.
Description of drawings
Fig. 1 is a Bump intake duct import wave system design diagram.
Fig. 2 is the Bump intake duct import wave system that obtains by flow average computation Mach number and total pressure recovery coefficient figure behind circular cone shock wave and the normal shock wave when different semi-cone angle.
Fig. 3 (a) is with schematic representation in the xoy coordinate surface of plane truncated cone shape shock wave generation rider body bulge compressing surface in supersonic flow.
Fig. 3 (b) is with schematic representation in the yoz coordinate surface of plane truncated cone shape shock wave generation rider body bulge compressing surface in supersonic flow.
Fig. 3 (c) is with schematic representation in the xoz coordinate surface of plane truncated cone shape shock wave generation rider body bulge compressing surface in supersonic flow.
Fig. 3 (d) is the rider body bulge compressing surface profile schematic three dimensional views that generates with plane truncated cone shape shock wave in supersonic flow.
Fig. 4 is a bulge compressing surface deflection angle theta schematic representation.
Fig. 5 is that different distance d place cuts semi-cone angle δ
cThe bulge compressing surface deflection angle theta of=12 ° the conical shock wave flow field gained that circular cone generated and the relation of d.
Fig. 6 is the relation apart from the ratio W/d of d and bulge width W and bulge compressing surface deflection angle theta.
Fig. 7 is a rider build face similar Design principle schematic.
Fig. 8 is bulge and a boundary layer thickness schematic representation in the intake duct import cross section.
Fig. 9 be import cross section bulge get rid of the boundary layer area than with bulge/every road height relationships figure.
Figure 10 is conformal import of intake duct import cross section and conical shock wave schematic representation.
Figure 11 (a) is the inlet total pres sure recovery coefficient comparison diagram of different lips sweepback angle scheme.
Figure 11 (b) is the intake duct resistance coefficient comparison diagram of different lips sweepback angle scheme.
Figure 12 is that the lip sweepback angle is 20 ° intake duct moulding figure.
Among the figure: 1, ultrasound velocity incoming flow, 2, bulge compressing surface, 3, conical shock wave, 4, inlet lip, 5, normal shock wave, 6, cone, 7, the plane of truncated cone shape shock wave, 8, the intersection (being rider body costa) of conical shock wave and plane taken, 9, bulge cross section, 10, fuselage molded lines, 11, the import boundary layer/every the road height and position, 12 inlet lip costas.
Embodiment
The present invention will contrast accompanying drawing below and give more fully to illustrate, given among each figure is an application example of the present invention, only is not confined to application example described herein and should not be construed to the present invention.
(1) do not wait the design of high-amplitude wave system
Fig. 1 illustrates one and adopts Bump intake duct import wave system schematic representation of the present invention.Ultrasound velocity incoming flow 1 produces conical shock wave 3 one at the head of bulge compressing surface 2, forms one normal shock wave 5 before inlet lip 4.Ma=1.6 is an example with design incoming flow Mach number, adopts the two wave systems design of " conical shock wave+normal shock wave ", and the total pressure recovery coefficient of intake duct multishock is σ
s=σ
1σ
2, σ wherein
1, σ
2Be respectively the total pressure recovery coefficient of first and second road shock wave, by waiting ripple to join the ripple theory analysis by force, total pressure recovery coefficient was the highest when the ripple of twice ripple equated by force, was best wave system.
Because behind the circular cone shock wave is conical flow, the flow field parameter skewness be to wait parameter line along the ray of crossing conical tip, so the average Mach number behind the circular cone shock wave generally calculates with the mean value of minimum and maximum Mach number between the conical surface and the shock surface.Fig. 2 has provided Mach number and total pressure recovery coefficient figure, wherein M α behind the circular cone shock wave of each the semi-cone angle correspondence that obtains by the flow average computation
1, M α
2Be respectively the Mach number behind first and second road shock wave, circular cone semi-cone angle δ
cIn the time of=24 °, the highest by total pressure recovery coefficient behind the ripple that waits the design of high-amplitude wave system, reach 0.985, corresponding shock wave angle is β
Awl=49.9 °, this is best wave angle β
Opt
Deng high-amplitude wave is that design is based on no viscosity flow, two dimensional surface shock theory design intake duct wave system.What the present invention adopted does not wait high-amplitude wave system to design, and is based on the consideration of following several respects:
1. the viscosity of fluid
Because the influence of fluid viscosity, meeting forms the boundary layer that development gradually thickens at solid wall surface, and the actual deflection angle of air-flow is increased, and wave system enhancing, shock wave angle are increased.
2. the heterogeneity in flow field
The heterogeneity in conical flow flow field, the shock wave angle when making the total pressure recovery coefficient maximum is less than best wave angle, and semi-cone angle is about 1 ° less than normal accordingly.
3. the low-speed performance of intake duct
Consider low-speed performance, the choosing of the semi-cone angle of circular cone should make shock wave lift-off not under each Mach number as far as possible, and therefore, the semi-cone angle of circular cone can not be excessive.Semi-cone angle is more little, and corresponding lift-off Mach number is just low more, yet it is big to depart from best wave system at this moment, and the wave system loss is also big, so semi-cone angle can not be too little, the span of suggestion semi-cone angle is 16 °＜δ
c＜25 °.
4. import/venturi Mach number requirement
For inferior, supersonic speed air inlet road, fitness for purpose inlet throat Mach number M α
t＜0.85, otherwise can cause the interior conduit loss to increase because airspeed is excessive.If wait when joining the design of ripple principle by force, then the conical shock wave ripple is bigger than normal by force, and Mach number is less than normal behind the ripple, (intake duct import) Mach number is bigger than normal after causing normal shock wave, air-flow is when passing through import to this section of venturi constricted channel, and Mach number can further increase, and causes the venturi Mach number excessive.In example, intake duct design Mach number 1.6 is a principle when carrying out the wave system design by waiting high-amplitude wave, and (intake duct import) Mach number promptly reaches 0.836～0.855 behind the normal shock wave, the venturi Mach number also will continue to increase, therefore, need to reduce inflow Mach number, must adopt and not wait the design of high-amplitude wave system.
To the throat area contraction ratio, calculating can get inflow Mach number and be not more than 0.75 and can meet the demands according to import, therefore chooses semi-cone angle δ
c=20 °, this moment circular cone shock wave angle β
Awl=45.8 °.
(2) based on the rider body bulge compressing surface similar Design method of profile drift angle
Determined circular cone semi-cone angle δ
c, Bump intake duct bulge compressing surface can obtain profile according to the generation body method of finding the solution conical flow.Fig. 3 shows the schematic representation of truncated cone shape shock wave generation bulge compressing surface in supersonic flow, and semi-cone angle is δ
c Circular cone 6 in supersonic flow, produce the conical shock wave 3 that semi-cone angle is β, use plane 7 truncated cone shape shock waves 3 apart from axial line distance d, the streamline that every bit sends backward on plane taken 7 and the conical shock wave intersection 8 has just constituted bulge compressing surface 2.
All do not provide the selection principle of plane taken 7 and circular cone 6 axial line distance d at present both at home and abroad about the design of rider body, the present invention has proposed a kind of similar Design method based on the profile angle of yaw according to the feature of conical flow.As shown in Figure 4, the profile deflection angle theta is defined as the tangent line of bulge compressing surface 2 vertical symmetry plane molded lines ends and intake duct the place ahead and comes angle between the flow path direction, because profile is a stream interface after the conical flow, so the profile angle of yaw is the terminal angle of yaw of streamline.
Fig. 5 shows section semi-cone angle δ at different distance d place
cThe bulge compressing surface bias angle theta of=12 ° the conical shock wave flow field gained that circular cone generated and the relation of d.The bulge compressing surface bias angle theta that different distance d place intercepts is different, along with the increase of d, δ
cReduce.Fig. 6 shows the relation apart from the ratio W/d of d and bulge width W and bulge compressing surface bias angle theta, different circular cone semi-cone angle, and W/d and profile deflection angle theta meet same Changing Pattern, and promptly when W/d 〉=10, the profile deflection angle theta approaches circular cone semi-cone angle δ
c, δ
c-θ＜1 °.
Fig. 7 shows rider build face similar Design principle schematic.Used two planar interception conical shock waves 3 of some A and some C to generate the rider body respectively, the length of rider body is respectively AB and CD, though the two apart from the circular cone axis apart from the d difference, but it is on the ray of φ that some B and some D all were positioned at circular cone 6 summit angles, face end is identical with the angle of circular cone 6 summit lines and circular cone axis as long as institute dams, the rider build face that is generated is just similar, and d is irrelevant with the intercepting height, and latter two profile of nondimensionalization overlaps fully.
Therefore, characteristic of the present invention is that design rider body no longer needs comparison and the optimization to parameter d, but the high-amplitude wave that do not wait that utilizes the present invention to propose is design, at first according to import wave system designing and calculating circular cone shock wave angle β
AwlWith circular cone semi-cone angle δ
c, determine the profile deflection angle theta then, generate the bulge compressing surface according to rider build face similar Design principle at last.
(3) set up relation between bulge design height and the local boundary layer thickness
One of main effect of bulge is to get rid of the fuselage boundary layer, owing to do not have boundary layer every the road, therefore bulge some be immersed in the fuselage boundary layer, so the design of bulge must be considered the relation with the fuselage boundary layer, i.e. relation between bump height h and the local boundary layer thickness δ.
Fig. 8 shows bulge and boundary layer thickness schematic representation in the intake duct import cross section, and bulge cross section 2 is with shadow representation among the figure.On fuselage 10 surfaces, one deck boundary layer is arranged, 11 expressions (dotted line 11 positions also are commonly used for the position of boundary layer every the road) of boundary layer thickness position with dashed lines, if boundary layer thickness is δ, ratio by given bump height h and δ, can be by the scaling of the given bulge compressing surface of h/ δ, the drop shadow curves of bulge in cross section of different sizes with every road and import antelabium line different intersection points is arranged.Suppose that the height after the preceding boundary layer of import is through the bulge surface is constant, still be δ, then because the effect of bulge the incoming flow boundary layer is moved to both sides row, so boundary layer exists only in two class delta-shaped regions of below, dotted line 11 positions, bulge top and antelabium line inboard.Use A
DiverterExpression is every the area of contour in road, A
BumpThe area of contour of expression bulge, A
FlowbyExpression correspondence boundary layer in the road area is excluded the area of part, promptly uses A
DiverterDeduct the remaining area of above-mentioned two class triangle areas, the approximate A that uses
Flwby/ A
DiverterThe eliminating effect of expression boundary layer.Work as A
Flowby/ A
DiverterRepresented that boundary layer was all got rid of at=1 o'clock.
Fig. 9 has provided import cross section bulge eliminating boundary layer area and has compared A
Flowby/ A
Diverter, bulge with compare A every the road area
Bump/ A
DiverterAnd bulge and the relation between road aspect ratio h/ δ.As can be seen, along with h/ δ is increased to 4 from 1, the eliminating amount of boundary layer increases, but the amplitude that increases slows down h/ δ=4 o'clock A gradually
Floby/ A
Diverte=0.929, promptly bump height is big more, and the boundary layer of eliminating is many more.
But on the other hand, bump height is big more, and then the wind-exposuring area of bulge is also big more.As can be seen, h/ δ=2 o'clock, the area of contour of bulge has approached the area of contour every the road, A
Bump/ A
Diverter=0.933.H/ δ=4 o'clock, the area of contour of bulge is 3.733 times every the road area of contour.The bulge area increases, and for satisfying the requirement of intake duct circulation area, then must raise the lip cover of intake duct, has increased the wind-exposuring area of whole intake duct thus again, and resistance also can correspondingly increase.Therefore, bump height can not be too little, and is highly too little, and then bulge major part buries in boundary layer, and the amount of getting rid of boundary layer is limited; Bump height can not be too big, and is highly too big, and then resistance is too big.
Bump height is not an independent parameter, restricted by a lot of design parameters.As design Mach number, generation body semi-cone angle δ
c, after parameter such as profile deflection angle theta determines, the profile of bulge has determined that its length and width height is all given, can need amplify the bulge compressing surface that is generated or dwindle according to actual size.Analyze from Fig. 9, h/ δ=2～2.5 are comparatively suitable, boundary layer eliminating this moment amount 0.793～0.842, and the bulge area of contour is 0.933～1.458 times every the road area of contour.
(4) conformal, the design of sweepback lip
Figure 10 shows conformal import and conical shock wave schematic representation in the intake duct import cross section.Lead general plane or the curved surface intercepting circular cone shock wave of adopting of rider body design based on the awl of conical shock wave and generate the bulge compressing surface, therefore have the mismatch problem of bulge compressing surface and fuselage surface.The present invention has considered the integrated requirement of fuselage when design, adopt fuselage 10 curved surfaces intercepting conical shock wave 3 to generate bulge compressing surface 2, so that bulge compressing surface 2 merges with fuselage 10 surface smoothing ground.
Because bulge 2 produces conical shock wave 3, and has formed a crescent shape zone between the fuselage 10, have only and when actinal surface shape and conical shock wave 3 are fitted fully, just can not cause overflow above the lip cover.As can see from Figure 10, owing to adopted the conformal import, make the most of and conical shock wave 3 of import antelabium fit, only there is overflow both sides such as intersecting with fuselage, and the overflow of both sides is inevitably, and the boundary layer that the bulge surface is excluded is flowed outward by two side direction imports just.
The Bump intake duct is got rid of the effect of boundary layer and is finished by bulge compressing surface and lip acting in conjunction.Figure 11 has provided the inlet characteristic comparison diagram of different lips sweepback angle scheme, and abscissa is inlet duct flow flow coefficient φ among the figure, and y coordinate is respectively total pressure recovery coefficient σ and resistance coefficient C
DAs can be seen, inlet total pres sure recovery coefficient curve and the parallel distribution of resistance coefficient curve approximation under the different sweepback angle, increase along with the lip sweepback angle, the total pressure recovery coefficient curve of intake duct rises, reach the highest when being 20 ° at the sweepback angle, the total pressure recovery coefficient curve descends again in the time of 30 °, between 10 ° and 20 °; The resistance coefficient curve then is to reduce along with the increase at lip sweepback angle, resistance coefficient minimum when the sweepback angle is 20 °, and curve is between 10 ° and 20 ° in the time of 30 °.It is 20 ° intake duct moulding figure that Figure 12 has provided the lip sweepback angle.
According to computational analysis, proposed the inlet lip sweepback angle should with the maximum deflection angle of bulge compressing surface, the corresponding to design principle of import normal shock wave angle.In the design process of Bump intake duct, the design of bulge compressing surface is relevant with the design of lip, can not be with the isolated design of these two-part.
The foregoing description just is used for explanation of the invention, and can not be as limitation of the present invention.Bump intake duct design Mach 2 ship 1.6 in this example, the present invention is applicable to that also all incoming flow Mach numbers are less than 2.0 Bump intake duct.Of the present inventionly do not wait high-amplitude wave system design, be applicable to the supersonic speed air inlet road of other type yet, comprise different entry shapes such as binary, axisymmetric, and external-compression type, mixing compression type wave system intake duct.Rider body bulge compressing surface similar Design method based on the profile drift angle of the present invention also is applicable to any ultrasound velocity, the design of hypersonic rider body, and the rider body generates and carries can be normal cone, also can be elliptic cone or other broad sense circular cone.Conformal of the present invention, sweepback lip design, and also are applicable to ultrasound velocity, the hypersonic inlet of other type.Therefore the mode of execution that mentality of designing every and of the present invention is identical is all in protection scope of the present invention.
Claims (5)
1. a realization does not wait high-amplitude wave to be and the method for the integrated Bump intake duct of forebody, and it is characterized in that: ultrasound velocity incoming flow (1) produces one conical shock wave (3) at the head of bulge compressing surface (2), at one normal shock wave of the preceding formation of inlet lip (4) (5);
The first step: the intake duct wave system adopts based on the external-compression type two wave system structures that do not wait high-amplitude wave system, and the total pressure recovery coefficient of intake duct multishock is σ
s=σ
1σ
2, σ wherein
1, σ
2Be respectively the total pressure recovery coefficient of conical shock wave (3), normal shock wave (5), by waiting ripple to join the ripple theory analysis by force, total pressure recovery coefficient was the highest when the ripple of twice ripple equated by force, was best wave system;
Second step: semi-cone angle is δ
cCircular cone (6) in supersonic flow, produce the conical shock wave (3) that semi-cone angle is β, the circle radius of conical shock wave (3) is R, plane (7) truncated cone shape shock wave (3) with distance cone axis linear distance d, wherein d＜R goes up the streamline formation bulge compressing surface that every bit sends backward from plane taken (7) and conical shock wave intersection (8);
The 3rd step: make that intake duct import bump height is h, local boundary layer thickness is δ, satisfy relation h/ δ=2～2.5 between intake duct import bump height h and the local boundary layer thickness δ, the bulge compressing surface that second step was generated carries out convergent-divergent, satisfies the actual size requirement;
The 4th step: inlet lip adopts conformal and the design of sweepback lip, the import antelabium is most of fits with the circular cone shock surface, the lip sweepback angle is consistent with the maximum deflection angle of bulge compressing surface, import normal shock wave angle respectively, to increase the total pressure recovery coefficient curve, to reduce resistance coefficient.
2. realization according to claim 1 does not wait the method for high-amplitude wave system and the integrated Bump intake duct of forebody, and it is characterized in that: the inflow Mach number behind the normal shock wave is not more than 0.75.
3. realization according to claim 1 does not wait the method for high-amplitude wave system and the integrated Bump intake duct of forebody, it is characterized in that: the plane or the curved surface that generate body circular cone axis different heights d with distance go to intercept the conical flow flow field, face is terminal identical with conical tip line and axis angle as long as institute dams, and the rider build face that is then generated is similar.
4. realization according to claim 1 does not wait the method for high-amplitude wave system and the integrated Bump intake duct of forebody, it is characterized in that: the import wave system according to intake duct calculates circular cone shock wave angle β
AwlWith circular cone semi-cone angle δ
c, determine the profile deflection angle theta then, generate the bulge compressing surface according to rider build face similar Design principle at last.
5. do not wait the method for high-amplitude wave system and the integrated Bump intake duct of forebody according to claim 1 or 4 described realizations, it is characterized in that: make that the bulge width is W, bulge width W and meet following Changing Pattern apart from ratio W/d between the d and bulge compressing surface bias angle theta, promptly when W/d 〉=10, the profile deflection angle theta approaches circular cone semi-cone angle δ
c, δ
c-θ＜1 °.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201010134882 CN101813027B (en) | 2010-03-29 | 2010-03-29 | Bump air inlet method for realizing integration of unequal-strength wave system with forebody |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201010134882 CN101813027B (en) | 2010-03-29 | 2010-03-29 | Bump air inlet method for realizing integration of unequal-strength wave system with forebody |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN101813027A true CN101813027A (en) | 2010-08-25 |
| CN101813027B CN101813027B (en) | 2013-04-10 |
Family
ID=42620373
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 201010134882 Expired - Fee Related CN101813027B (en) | 2010-03-29 | 2010-03-29 | Bump air inlet method for realizing integration of unequal-strength wave system with forebody |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN101813027B (en) |
Cited By (20)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102518517A (en) * | 2011-12-08 | 2012-06-27 | 南京航空航天大学 | Bistable air inlet, its application as aircraft air inlet and design method of bistable air inlet |
| CN102828832A (en) * | 2012-08-14 | 2012-12-19 | 西北工业大学 | Method for improving starting capability of fixed-geometry two-dimensional mixed-compression inlet |
| CN102996253A (en) * | 2012-12-31 | 2013-03-27 | 中国人民解放军国防科学技术大学 | Supersonic air intake duct and wall face determination method of supersonic air intake duct |
| CN103225542A (en) * | 2013-04-11 | 2013-07-31 | 南京航空航天大学 | Deformable bulging gas inlet channel bulging molded surface deformation achievement method |
| CN103963996A (en) * | 2014-05-23 | 2014-08-06 | 厦门大学 | Transverse pressure gradient controlled ride wave forebody and air inlet channel integrated designing method |
| CN104384288A (en) * | 2014-11-19 | 2015-03-04 | 中国航空工业集团公司沈阳飞机设计研究所 | Adaptive bump air inlet passage shape control method based on flexible skin |
| CN105697150A (en) * | 2016-03-07 | 2016-06-22 | 厦门大学 | Bump air inlet design method integrating cone-derived and osculating waveriders |
| CN106218865A (en) * | 2016-09-13 | 2016-12-14 | 中国航空工业集团公司沈阳空气动力研究所 | A kind of aircraft precursor eddy current dynamic control device |
| CN106438047A (en) * | 2015-08-04 | 2017-02-22 | 北京机电工程研究所 | Buried gas inlet channel inner channel design method |
| CN106640378A (en) * | 2016-11-24 | 2017-05-10 | 中国航空工业集团公司沈阳飞机设计研究所 | Air inlet channel main bearing structure and air inlet channel main bearing structure deformation control method |
| CN107215473A (en) * | 2017-06-08 | 2017-09-29 | 南京航空航天大学 | A kind of nothing integrated with aircraft is every road Subsonic inlet |
| CN109720586A (en) * | 2017-10-30 | 2019-05-07 | 成都飞机工业（集团）有限责任公司 | A kind of boundary layer diverter |
| CN109899178A (en) * | 2019-03-08 | 2019-06-18 | 中国人民解放军国防科技大学 | Hypersonic air inlet channel with pre-compression device |
| CN110990955A (en) * | 2019-12-12 | 2020-04-10 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and system |
| CN111942600A (en) * | 2020-08-06 | 2020-11-17 | 四川航天中天动力装备有限责任公司 | Boundary layer-free partition air inlet channel |
| CN112224425A (en) * | 2020-09-09 | 2021-01-15 | 西北工业大学 | Conformal design method of aircraft air inlet structure and aircraft |
| CN112298598A (en) * | 2020-11-02 | 2021-02-02 | 厦门大学 | Hypersonic bulge compression profile reverse design method based on curved cone precursor |
| CN112340014A (en) * | 2020-11-06 | 2021-02-09 | 南京航空航天大学 | Inner-outer flow decoupling double-waverider high-speed air suction type aircraft and generation method thereof |
| CN112985755A (en) * | 2021-05-20 | 2021-06-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Boundary layer similar parameter simulation method for accurately predicting cavity flow acoustic load |
| CN115653754A (en) * | 2022-12-12 | 2023-01-31 | 中国航空工业集团公司西安飞机设计研究所 | Supersonic air inlet system with three wave systems for fixing compression surface |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4611616A (en) * | 1984-01-10 | 1986-09-16 | Messerschmitt-Bolkow-Blohm Gmbh | Axially semisymmetrical supersonic air intake for reaction engines, particularly solid fuel ram jet rocket engines |
| CN101016847A (en) * | 2007-02-27 | 2007-08-15 | 南京航空航天大学 | High supersound air-intake air turbogenerator |
| CN101392685A (en) * | 2008-10-29 | 2009-03-25 | 南京航空航天大学 | Internal waverider hypersonic inlet and design method based on random shock form |
| CN101418723A (en) * | 2008-10-15 | 2009-04-29 | 南京航空航天大学 | Internal waverider-derived hypersonic inlet with ordered inlet and outlet shape and design method |
-
2010
- 2010-03-29 CN CN 201010134882 patent/CN101813027B/en not_active Expired - Fee Related
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4611616A (en) * | 1984-01-10 | 1986-09-16 | Messerschmitt-Bolkow-Blohm Gmbh | Axially semisymmetrical supersonic air intake for reaction engines, particularly solid fuel ram jet rocket engines |
| CN101016847A (en) * | 2007-02-27 | 2007-08-15 | 南京航空航天大学 | High supersound air-intake air turbogenerator |
| CN101418723A (en) * | 2008-10-15 | 2009-04-29 | 南京航空航天大学 | Internal waverider-derived hypersonic inlet with ordered inlet and outlet shape and design method |
| CN101392685A (en) * | 2008-10-29 | 2009-03-25 | 南京航空航天大学 | Internal waverider hypersonic inlet and design method based on random shock form |
Cited By (32)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102518517A (en) * | 2011-12-08 | 2012-06-27 | 南京航空航天大学 | Bistable air inlet, its application as aircraft air inlet and design method of bistable air inlet |
| CN102518517B (en) * | 2011-12-08 | 2014-03-05 | 南京航空航天大学 | Design method of bistable air inlet |
| CN102828832A (en) * | 2012-08-14 | 2012-12-19 | 西北工业大学 | Method for improving starting capability of fixed-geometry two-dimensional mixed-compression inlet |
| CN102828832B (en) * | 2012-08-14 | 2014-05-14 | 西北工业大学 | Method for improving starting capability of fixed-geometry two-dimensional mixed-compression inlet |
| CN102996253A (en) * | 2012-12-31 | 2013-03-27 | 中国人民解放军国防科学技术大学 | Supersonic air intake duct and wall face determination method of supersonic air intake duct |
| CN102996253B (en) * | 2012-12-31 | 2015-03-04 | 中国人民解放军国防科学技术大学 | Supersonic air intake duct and wall face determination method of supersonic air intake duct |
| CN103225542A (en) * | 2013-04-11 | 2013-07-31 | 南京航空航天大学 | Deformable bulging gas inlet channel bulging molded surface deformation achievement method |
| CN103225542B (en) * | 2013-04-11 | 2015-06-03 | 南京航空航天大学 | Deformable bulging gas inlet channel bulging molded surface deformation achievement method |
| CN103963996A (en) * | 2014-05-23 | 2014-08-06 | 厦门大学 | Transverse pressure gradient controlled ride wave forebody and air inlet channel integrated designing method |
| CN103963996B (en) * | 2014-05-23 | 2015-11-11 | 厦门大学 | The waverider forebody derived that transverse-pressure gradient is controlled and inlet channel integrated design method |
| CN104384288A (en) * | 2014-11-19 | 2015-03-04 | 中国航空工业集团公司沈阳飞机设计研究所 | Adaptive bump air inlet passage shape control method based on flexible skin |
| CN106438047B (en) * | 2015-08-04 | 2018-04-10 | 北京机电工程研究所 | The design method of Submerged Inlet internal channel |
| CN106438047A (en) * | 2015-08-04 | 2017-02-22 | 北京机电工程研究所 | Buried gas inlet channel inner channel design method |
| CN105697150A (en) * | 2016-03-07 | 2016-06-22 | 厦门大学 | Bump air inlet design method integrating cone-derived and osculating waveriders |
| CN106218865A (en) * | 2016-09-13 | 2016-12-14 | 中国航空工业集团公司沈阳空气动力研究所 | A kind of aircraft precursor eddy current dynamic control device |
| CN106640378A (en) * | 2016-11-24 | 2017-05-10 | 中国航空工业集团公司沈阳飞机设计研究所 | Air inlet channel main bearing structure and air inlet channel main bearing structure deformation control method |
| CN106640378B (en) * | 2016-11-24 | 2018-06-29 | 中国航空工业集团公司沈阳飞机设计研究所 | A kind of main bearing structure of air intake duct and the main bearing structure deformation control method of air intake duct |
| CN107215473A (en) * | 2017-06-08 | 2017-09-29 | 南京航空航天大学 | A kind of nothing integrated with aircraft is every road Subsonic inlet |
| CN107215473B (en) * | 2017-06-08 | 2018-08-31 | 南京航空航天大学 | A kind of and integrated nothing of aircraft is every road Subsonic inlet |
| CN109720586A (en) * | 2017-10-30 | 2019-05-07 | 成都飞机工业（集团）有限责任公司 | A kind of boundary layer diverter |
| CN109899178A (en) * | 2019-03-08 | 2019-06-18 | 中国人民解放军国防科技大学 | Hypersonic air inlet channel with pre-compression device |
| CN110990955B (en) * | 2019-12-12 | 2024-05-28 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system |
| CN110990955A (en) * | 2019-12-12 | 2020-04-10 | 中国人民解放军国防科技大学 | Hypersonic speed Bump air inlet channel design method and system |
| CN111942600A (en) * | 2020-08-06 | 2020-11-17 | 四川航天中天动力装备有限责任公司 | Boundary layer-free partition air inlet channel |
| CN112224425A (en) * | 2020-09-09 | 2021-01-15 | 西北工业大学 | Conformal design method of aircraft air inlet structure and aircraft |
| CN112224425B (en) * | 2020-09-09 | 2024-06-07 | 西北工业大学 | Shape-preserving design method for air inlet channel structure of aircraft and aircraft |
| CN112298598A (en) * | 2020-11-02 | 2021-02-02 | 厦门大学 | Hypersonic bulge compression profile reverse design method based on curved cone precursor |
| CN112340014A (en) * | 2020-11-06 | 2021-02-09 | 南京航空航天大学 | Inner-outer flow decoupling double-waverider high-speed air suction type aircraft and generation method thereof |
| WO2022095163A1 (en) * | 2020-11-06 | 2022-05-12 | 南京航空航天大学 | Internal-external flow decoupled dual-waverider high-speed airbreathing aircraft and generation method therefor |
| CN112340014B (en) * | 2020-11-06 | 2022-05-13 | 南京航空航天大学 | Inner-outer flow decoupling double-waverider high-speed air suction type aircraft and generation method thereof |
| CN112985755A (en) * | 2021-05-20 | 2021-06-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Boundary layer similar parameter simulation method for accurately predicting cavity flow acoustic load |
| CN115653754A (en) * | 2022-12-12 | 2023-01-31 | 中国航空工业集团公司西安飞机设计研究所 | Supersonic air inlet system with three wave systems for fixing compression surface |
Also Published As
| Publication number | Publication date |
|---|---|
| CN101813027B (en) | 2013-04-10 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN101813027B (en) | Bump air inlet method for realizing integration of unequal-strength wave system with forebody | |
| CN112340014B (en) | Inner-outer flow decoupling double-waverider high-speed air suction type aircraft and generation method thereof | |
| CN101798961B (en) | Two-stage beveled supersonic speed air inlet lip | |
| CN104908957A (en) | Ridge type sweeping vortex generator and generation method | |
| CN106394874B (en) | A kind of Waverider of double sweepback layouts | |
| CN212337458U (en) | Three-dimensional inner rotation damping expansion wave high-performance double-channel TBCC air inlet | |
| CN106741976A (en) | A kind of mimetic design method of waverider forebody derived air intake duct integration configuration | |
| KR102616901B1 (en) | Aerodynamic layout design method and system for wide-speed-range hypersonic aircraft | |
| CN113279860B (en) | Inner bulge S-shaped air inlet channel with middle control section and method | |
| CN110990955B (en) | Hypersonic speed Bump air inlet channel design method and hypersonic speed Bump air inlet channel design system | |
| CN108038295A (en) | Hypersonic inlet channel and isolation section integrated design method | |
| CN115056998A (en) | Longitudinal segmentation and hierarchical compression design method for cone guide multiplier front body | |
| CN114802776A (en) | Embedded air inlet channel based on sweepback step displacement precursor boundary layer | |
| CN117902051B (en) | Air inlet channel adopting embedded micro-channel array and design method thereof | |
| CN108412618A (en) | Hypersonic/supersonic axisymmetric inlet lip and design method thereof | |
| CN113104198B (en) | Wave-rider precursor self-adaptive boundary layer displacement inward-transfer air inlet channel integrated design method | |
| CN105539863A (en) | Integrated aerodynamic layout method for hypersonic aircraft forebody, air inlet duct and supporting plate | |
| CN112298598B (en) | Hypersonic bulge compression profile reverse design method based on curved cone precursor | |
| CN109723571B (en) | Throat offset type pneumatic thrust vectoring nozzle with trapezoidal section and aircraft equipped with same | |
| CN111942600B (en) | Boundary layer-free partition air inlet channel | |
| CN201301753Y (en) | Inner wave rider type air inlet channel taking internal and external flow performance into consideration | |
| CN103029830A (en) | Double waverider splicing air suction type hypersonic aircraft forebody and design method thereof | |
| CN114261530B (en) | Integrated design method for minimum resistance cone-guided wave-rider and three-dimensional inward-rotation type air inlet channel | |
| CN114802799B (en) | Full three-dimensional two-stage compression double waverider integrated design method based on bending shock wave theory | |
| CN214035885U (en) | Three-dimensional inward-rotation air inlet channel with curved drainage tube |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20130410 Termination date: 20200329 |
|
| CF01 | Termination of patent right due to non-payment of annual fee |