CN105151307B - Method for cutting Mach surface of hypersonic aircraft with forebody/air inlet pipeline in integrated design - Google Patents
Method for cutting Mach surface of hypersonic aircraft with forebody/air inlet pipeline in integrated design Download PDFInfo
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- CN105151307B CN105151307B CN201510645281.0A CN201510645281A CN105151307B CN 105151307 B CN105151307 B CN 105151307B CN 201510645281 A CN201510645281 A CN 201510645281A CN 105151307 B CN105151307 B CN 105151307B
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- inlet pipeline
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Abstract
The invention discloses a method for cutting a Mach surface of a hypersonic aircraft with a forebody/air inlet pipeline in integrated design, and belongs to the technical field of design of hypersonic aircrafts. The method comprises the following steps: designing the ratio of the width to the height of a section of the air inlet pipeline according to the requirements of a combustion chamber; according to design parameters of a compression system of the air inlet pipeline, determining parameters of the incoming flow of the air inlet pipeline and a Mach number, a shock wave angle and airflow turning angle parameters of each level of compression surfaces; according to the width of an inlet of the air inlet pipeline and the Mach number of the compression surface superior to the air inlet pipeline, tracing a Mach line upstream so as to form the edge of each level of the compression surfaces; and then according to the shock wave angle of each level of the compression surfaces, obtaining the final compression system of external compression/air inlet pipeline. Through the adoption of the method for cutting the Mach surface, disclosed by the invention, uniform air currents of the air inlet pipeline can be guaranteed through the obtained forebody, besides the width of the inlet of the air inlet pipeline is equal to the width of the air inlet pipeline, and the situation that the shock wave drag is increased because the width of the forebody is increased is avoided, so that the lift-drag ratio of the whole aircraft is increased.
Description
Technical field
The invention belongs to hypersonic aircraft design field and in particular to a kind of hypersonic aircraft precursor/
The mach front cutting method of air intake duct integrated design.
Background technology
Forebody/Inlet integrated design is one of key technology of Air-breathing hypersonic vehicle design.Under precursor
Surface is equivalent to the external compression face of scramjet engine, air-flow Mach number needed for speed is down to air intake duct after shock wave compression,
Flow into air intake duct again to be further compressed, be finally reached Mach number needed for burning, temperature and pressure.Therefore, Forebody/Inlet one
Bodyization design directly affects engine performance, simultaneously because precursor is also a part for body, precursor design equally can affect to fly
The aeroperformance of row device.
The integrated design method commonly used at present is using sphenoid as shock wave compression unit, by multistage external pressure be condensed to into
Air flue provides uniform incoming flow.However, under conditions of no side plate blocks, after ripple, high pressure gas are known from experience and are gradually leaked to both sides, lead to
The problem of inlet mouth marginal existence imperfect flow.For avoiding this problem, a solution is so that precursor width is more than
Air intake duct width, to reduce the gas leakage of precursor both sides to entering the impact of inlet flow it is ensured that air intake duct carrys out family status matter, such as
Shown in Fig. 1 and Fig. 2 two kind aircraft air intake duct.However, because wedge shape precursor needs to provide the Mach number needed for air intake duct, entering
Before air flue lip the object plane angle of compressing surface often than larger so that compressing surface drag due to shock wave is larger.Now increase precursor again
Width makes aircraft front face area increase again, further increases resistance, reduces vehicle lift-drag.
Content of the invention
In order to solve problems of the prior art, the present invention proposes a kind of hypersonic aircraft Forebody/Inlet
The mach front cutting method of integrated design.According to theory of characteristics, for a sphenoid shock wave flow field, if before with sphenoid
Edge angle point is summit, and along after wedge shape bulk wave, sphenoid is subject to the Partial Resection of side airflow expansion effects by mach front, can't
Change Mach cone with the flow field of exterior domain.According to this thought, for two-dimentional wedge shape precursor, using upstream tracking mach line
Method carries out Forebody/Inlet integrated design, can avoid increasing precursor width band ensureing inlet flow while uniform
The drag due to shock wave come increases.
From theory of characteristics, in two dimension no viscosity flow field, there is three characteristic curves, i.e. streamline and left and right rows mach line, stream
Field information dependence characteristics line is downstream propagated it is impossible to cross over characteristic curve propagation regions；In three-dimensional flow field information of flow not across
Characteristic face and the propagation regions of stream interface.The unbaffled sphenoid for both sides, from the small pertubation theory of supersonic flow field, swashs
In supersonic flow field after ripple, the spread scope of microvariations is Mach cone, therefore the coverage that no baffle plate build-up of pressure leaks is
Mach cone with sphenoid leading edge angle point as summit.Therefore, for sphenoid shock wave flow field, if with sphenoid leading edge angle point being
Summit, along after wedge shape bulk wave, sphenoid is subject to the Partial Resection of side airflow expansion effects by mach front, then will not change Mach
Cone is with the flow field of exterior domain.
The mach front cutting method of the hypersonic aircraft Forebody/Inlet integrated design that the present invention provides, including
Following steps:
Step 1: by the ratio of width to height of combustor Demand Design inlet channel cross-sectional.To using wedge shape precursor as external compression part
Two-dimentional air intake duct, combustor adopt square-section combustor.
Step 2: by the design parameter of air intake duct compressibility, determine air intake duct flowing parameter, and according to equal strength group
Knit wave system method and determine compressing surface Mach numbers at different levels, Angle of Shock Waves, air-flow turnover angular dimensions.
The design parameter of described air intake duct compressibility includes free stream Mach number, inlet mouth Mach number.
Step 3: by inlet mouth width and its higher level's compressing surface Mach number, upstream follow the trail of mach line, formed last
One stage of compression face edge.Again by afterbody compressing surface, mach line is upstream followed the trail of according to its higher level's compressing surface Mach number,
Form penultimate stage compressing surface edge.By that analogy, compressing surface edges at different levels are obtained.Further according to the pressure at different levels of gained in step 2
Contracting face Angle of Shock Waves, obtains final external compression/air intake duct compressibility.
It is an advantage of the current invention that:
Body and air intake duct coupling are often examined by traditional air suction type hypersonic Waverider aircraft integrated design method
Consider, that is, wish body and air intake duct in the shock wave flow field of same intensity, horizontal at inlet mouth to reduce
Flowing.But the design object of air intake duct and body is different.Air intake duct needs by the larger object plane of the angle of attack to flowing to
Row compression, just can make hypersonic air-flow be reduced to the Mach number of combustor needs.And body object plane then needs to meet less
Could produce under angle compared with high lift-drag ratio.Therefore if it is considered that body is identical with the shock strength of air intake duct, couple body and entered
Air flue design, necessarily causes that body object plane angle is larger, and lift-drag ratio is less.The most important meaning of mach front cutting method is can be one
Determine in degree to decouple the design of rider body and air intake duct.Because the object plane angle that body produces during high lift-drag ratio is typically less than
It is equal to the object plane angle in integrated air intake duct precommpression face, so air intake duct is after following the trail of generation compressing surface based on mach line, it sets
Meter target is to consider engine performance parameter and air intake duct startability；And the design object of rider body is then to pursue promotion
Resistance ratio, the lift-drag ratio loss causing for coupling punching engine leaves sufficient space.Understand according to the above analysis, due to air inlet
The shock wave that road produces is better than body, therefore in order to ensure that following the trail of gases at high pressure after the air intake duct shock wave that mach line is formed does not leak
Block upper surface is it is desirable to body leading edge sweep is preferably more than the Angle of Shock Waves in air intake duct one stage of compression face.Using the present invention's
Mach front cutting method gained precursor can guarantee that inlet flow uniformly, simultaneously at inlet mouth with air intake duct width phase
Deng, it is to avoid increase the drag due to shock wave that precursor width causes and increase, thus improve full machine lift-drag ratio.
Brief description
Fig. 1 is Russian igla air intake duct schematic diagram；
Fig. 2 is U.S. hssw aircraft air intake duct schematic diagram；
Fig. 3 is sphenoid and the sphenoid flow field contrast along after mach front cutting, and in figure condition is: ma=5.0, δ=10 °；
Fig. 4 integrated Forebody/Inlet compressibility top view；
Fig. 5 integrated Forebody/Inlet compressibility perspective view；
Grade mach line cloud atlas on Fig. 6 integrated Forebody/Inlet plane of symmetry；
Grade mach line cloud atlas on Fig. 7 integrated Forebody/Inlet different cross section.
In figure:
1. one stage of compression face edge；2. two-stage compression face edge；3. inlet mouth；4. three stage compression face edge；5. four
Level compressing surface edge；6. level Four compressing surface leading edge；7. edge in face of three stage compression；8. edge in face of two-stage compression；9. one stage of compression face
Leading edge.
Specific embodiment
The present invention is described in detail for embodiment below in conjunction with the accompanying drawings.
The present invention proposes a kind of mach front cutting method of hypersonic aircraft Forebody/Inlet integrated design, pin
To two-dimentional wedge shape precursor, according to mach front incision principle, integrated precursor/enter is formed using the method upstream following the trail of mach line
Air flue, avoids increasing the drag due to shock wave increase that precursor width brings while guarantee inlet flow is uniform.
The mach front cutting method of the hypersonic aircraft Forebody/Inlet integrated design that the present invention provides, including
Following steps:
Step 1: by the ratio of width to height of combustor Demand Design inlet channel cross-sectional.To using wedge shape precursor as external compression part
Two-dimentional air intake duct, external compression face be rely on Two-Dimensional Shock be compressed, the contraction of flow tube all concentrates on short transverse, thus burning
Room can adopt square-section combustor.
Step 2: by the design parameter of air intake duct compressibility, determine air intake duct flowing parameter.
First, Mach number is designed according to flight Mach number and electromotor, determine wave system quantity.Then, organize by equal strength
Wave system method, determines compressing surface Mach numbers at different levels, Angle of Shock Waves, air-flow turnover angular dimensions, Shi Ge road oblique shock wave wavefront Mach number
Normal component equal it may be assumed that
ma1sinβ1=ma2sinβ2=...=man-2sinβn-2(1)
After shock wave front, Mach number relational expression is:
Wherein, mai、mai+1It is respectively wavefront, Mach number after ripple.
Shock wave angle beta with the relational expression of air-flow knuckle δ is:
In various above, maiFor Mach number, βiFor Angle of Shock Waves, γ is specific heats of gases ratio, and δ is air-flow knuckle, i=1,
2 ..., n, n are compressing surface series.
In the bar that the design parameter (including free stream Mach number and inlet mouth Mach number) of air intake duct compressibility is given
Under part, simultaneous above formula (1)～(3), Mach number after solution each road shock front, ripple, and compressing surface Mach number at different levels, shock wave
Angle, air-flow knuckle etc..
Step 3: according to compressing surface Mach numbers at different levels, by inlet mouth width and its higher level's compressing surface Mach number, upwards
Mach line is followed the trail of in trip, forms higher level's compressing surface edge.Specifically, if higher level is compressing surface Mach 2 ship ma2, then entered by air intake duct
Mouth marginal point starts, and makees Mach angle μ=arcsin (1/ma2) mach line, follow the trail of this mach line and form higher level's compressing surface side
Edge.So repeatedly, compressing surface edges at different levels are obtained.Further according to the compressing surface at different levels Angle of Shock Waves of gained in step 2, obtain final outer
Compression/air intake duct compressibility.
Fig. 3 gives the shock wave flow field of sphenoid and the contrast in former sphenoid shock wave flow field along after mach front cutting, can see
After going out excision, disturbance does not travel to the central area below sphenoid, and the flowing in central area does not have any change, explanation
Cutting flow field along mach front can't affect other regions in flow field.
Embodiment: design free stream Mach number ma=5, air intake port Mach number ma=2.The Mach being provided according to the present invention
The integrated Forebody/Inlet that face cutting method design obtains is as shown in Figure 4 and Figure 5.In this embodiment, inlet mouth 3 is one
Square-section entrance, designed compressing surface is level Four compressing surface, and wherein lip puts the third level and fourth stage compressing surface with lining, front
The body arrangement first order and second level compressing surface.
Lip with inner compressing surface method for designing is: by inlet mouth 3 lower edge, inverse direction of flow, that is, upstream
Follow the trail of level Four compressing surface edge 5, obtain fourth stage compressing surface, the position of level Four compressing surface leading edge 6, air-flow should be made to pass through this level Four
The shock wave producing after compressing surface leading edge 6 meets at inlet mouth 3 upper limb；Again three-level is upstream followed the trail of by level Four compressing surface leading edge 6
Compressing surface edge 4, obtains third level compressing surface, and in face of three stage compression, the position of edge 7 should make air-flow pass through in face of this three stage compression
The shock wave producing after edge 7 meets at inlet mouth 3 upper limb.
The compressing surface method for designing of precursor is: by inlet mouth 3 upper limb, upstream follows the trail of two-stage compression face edge
2, obtain second level compressing surface, in face of two-stage compression, the position of edge 8 should make air-flow produce after edge 8 in face of this two-stage compression
Shock wave meets at lip；Again one stage of compression face edge 1 is upstream followed the trail of by edge in face of two-stage compression 8, obtain first order compressing surface, one
The position of level compressing surface leading edge 9 should make the shock wave that air-flow produces after edge 9 in face of this one stage of compression meet at lip.
As seen from Figure 5, gained inlet lip position sweepforward, the compressing surface width upstream following the trail of formation along mach line is big
In inlet mouth width.External pressure compression system both sides need not add baffle plate, also can ensure the uniform incoming flow that air intake duct needs, simultaneously
Both sides can be as overflow ducts, to improve the starting performance of air intake duct.
Numerical simulation is carried out to designed air intake duct using euler equation, being calculated total pressure recovery coefficient is 0.8270.
Fig. 6 gives and waits mach line cloud atlas it can be seen that the twice shock wave of air intake duct upper wall surface on the integrated Forebody/Inlet plane of symmetry
All converge in lower wall surface lip position, the twice shock wave of lower wall surface all converges in inlet mouth, and after ripple, parameter uniformly, shows
Design Theory is coincide with number analog result.Fig. 7 gives second and third grade of compressing surface, and air intake duct leading edge, air intake port section
The grade mach line cloud atlas of euler equation.Can be seen that the precursor of application mach line tracer technique design can provide product for air intake duct
Of fine quality good flowing.
Claims (3)
1. the integrated design of hypersonic aircraft Forebody/Inlet mach front cutting method it is characterised in that:
Step one, by combustor Demand Design inlet channel cross-sectional the ratio of width to height, to using wedge shape precursor as external compression part
Two-dimentional air intake duct, combustor adopts square-section combustor；
Step 2, by air intake duct compressibility design parameter, determine air intake duct flowing parameter, and organize ripple according to equal strength
System, method determines compressing surface Mach numbers at different levels, Angle of Shock Waves, air-flow deflection angle；
The design parameter of described air intake duct compressibility includes free stream Mach number, inlet mouth Mach number；
Step 3, by inlet mouth width and described air intake duct higher level's compressing surface Mach number, upstream follow the trail of mach line, shape
Become afterbody compressing surface edge；Again by afterbody compressing surface, according to higher level's compression of described afterbody compressing surface
Face Mach number upstream follows the trail of mach line, forms penultimate stage compressing surface edge；By that analogy, compressing surface sides at different levels are obtained
Edge；Further according to the compressing surface at different levels Angle of Shock Waves of gained in step 2, obtain final external compression/air intake duct compressibility.
2. the mach front cutting method of hypersonic aircraft Forebody/Inlet integrated design according to claim 1,
It is characterized in that: the particular content of step 2 is:
First, Mach number is designed according to flight Mach number and electromotor, determine wave system quantity；
Then, organize wave system method by equal strength, determine compressing surface Mach numbers at different levels, Angle of Shock Waves, air-flow turnover angular dimensions, make each
The normal component of road oblique shock wave wavefront Mach number equal it may be assumed that
ma1sinβ1=ma2sinβ2=...=man-2sinβn-2(1)
After shock wave front, Mach number relational expression is:
Shock wave angle beta with the relational expression of air-flow deflection angle δ is:
In various above, mai、mai+1It is respectively the wavefront of i-stage compressing surface, Mach number, β after rippleiFor i-stage compressing surface
Angle of Shock Waves, γ is specific heats of gases ratio, and δ is air-flow deflection angle, i=1, and 2 ..., n, n are compressing surface series；
Under conditions of the design parameter of air intake duct compressibility gives, simultaneous above formula (1)～(3), solve each road shock wave ripple
Before, Mach number after ripple, and compressing surface Mach number at different levels, Angle of Shock Waves, air-flow deflection angle.
3. the mach front cutting method of hypersonic aircraft Forebody/Inlet integrated design according to claim 1,
It is characterized in that: the mach line described in step 3 refers to, if higher level is compressing surface Mach 2 ship ma2, then by inlet mouth side
Edge point starts, and makees Mach angle μ=arcsin (1/ma2) mach line.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110162901A (en) * | 2019-05-28 | 2019-08-23 | 中国人民解放军国防科技大学 | Optimized design method and system for axisymmetric configuration precursor of hypersonic aircraft |
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| CN105667811B (en) * | 2016-01-27 | 2017-11-07 | 南京航空航天大学 | The design method of hypersonic aircraft precursor and the multistage coupling integrated configuration of air intake duct |
| CN105539863B (en) * | 2016-01-29 | 2017-06-13 | 南京航空航天大学 | Hypersonic aircraft precursor, air intake duct and support plate integrated pneumatic layout method |
| CN105738067B (en) * | 2016-02-01 | 2018-04-06 | 南京航空航天大学 | The fast determination method of parameter after a kind of twice homonymy oblique shock wave is intersecting |
| CN106043737B (en) * | 2016-06-29 | 2018-06-12 | 中国人民解放军国防科学技术大学 | A kind of " waiting object planes-change Mach number " wide fast domain Waverider aircraft design method |
| CN106043738B (en) * | 2016-06-29 | 2018-08-28 | 中国人民解放军国防科学技术大学 | A kind of equal shock waves flow field-change Mach SerComm speed domain Waverider aircraft design method |
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