CN105151307A - 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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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, be specifically related to the mach front cutting method of a kind of hypersonic aircraft Forebody/Inlet integrated design.
Background technology
Forebody/Inlet integrated design is one of gordian technique of Air-breathing hypersonic vehicle design.Precursor lower surface is equivalent to the external compression face of scramjet engine, and air-flow speed after shock wave compression is down to Mach number needed for inlet channel, then flows into inlet channel and be further compressed, and finally reaches burning required Mach number, temperature and pressure.Therefore, Forebody/Inlet integrated design directly affects engine performance, and simultaneously due to the part that precursor is also body, precursor design can affect the aeroperformance of aircraft equally.
Integrated design method conventional is at present using sphenoid as shock wave compression unit, is condensed to inlet channel provides uniform incoming flow by multistage external pressure.But under the condition of blocking without side plate, after ripple, high pressure gas cognition leaks to both sides gradually, causes the problem of inlet mouth marginal existence imperfect flow.For avoiding this problem, a solution makes precursor width be greater than inlet channel width, to reduce precursor both sides Leakage Gas to the impact entering inlet flow, ensures inlet channel incoming flow quality, two kinds of aircraft inlet channels as depicted in figs. 1 and 2.But because wedge shape precursor needs the Mach number that provides needed for inlet channel, before inlet lip, the object plane angle of compressing surface is often larger, makes compressing surface drag due to shock wave larger.Now increasing precursor width again makes again aircraft wind area increase, and further increases resistance, reduces vehicle lift-drag.
Summary of the invention
In order to solve problems of the prior art, the present invention proposes the mach front cutting method of a kind of hypersonic aircraft Forebody/Inlet integrated design.According to theory of characteristics, for a sphenoid shock wave flow field, if with sphenoid leading edge angle point for summit, 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, adopt the method for upstream following the trail of mach line to carry out Forebody/Inlet integrated design, can ensure that the drag due to shock wave that inlet flow avoids increase precursor width to bring uniformly simultaneously increases.
From theory of characteristics, there is three characteristic curves, i.e. streamline and left and right rows mach line without in viscosity flow field in two dimension, information of flow dependence characteristics alignment downstream travel, can not cross over characteristic curve propagation regions; In three-dimensional flow field, information of flow can not cross over the propagation regions of characteristic face and stream interface.For the unbaffled sphenoid in both sides, from the small pertubation theory of supersonic flow field, in supersonic flow field after shock wave, the spread scope of microvariations is Mach cone, therefore is the Mach cone that is summit with sphenoid leading edge angle point without the incidence that baffle plate build-up of pressure leaks.Therefore, for sphenoid shock wave flow field, if with sphenoid leading edge angle point for summit, after wedge shape bulk wave, sphenoid is subject to the Partial Resection of side airflow expansion effects by mach front, then can not change Mach cone with the flow field of exterior domain.
The mach front cutting method of hypersonic aircraft Forebody/Inlet provided by the invention integrated design, comprises the following steps:
Step 1: by the ratio of width to height of combustion chamber Demand Design inlet channel cross-sectional.To the two-dimentional inlet channel adopting wedge shape precursor as external compression part, combustion chamber adopts square-section combustion chamber.
Step 2: by the design parameters of inlet channel compression system, determines the incoming flow parameter of inlet channel, and organizes wave system method to determine compressing surface Mach number at different levels, Angle of Shock Waves, air-flow turning angle parameter according to equal strength.
The design parameters of described inlet channel compression system comprises 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, forms afterbody compressing surface edge.Again by afterbody compressing surface, upstream follow the trail of mach line according to its higher level's compressing surface Mach number, form penultimate stage compressing surface edge.By that analogy, compressing surface edge at different levels is obtained.Again according to the compressing surface Angle of Shock Waves at different levels of gained in step 2, obtain final external compression/inlet channel compression system.
The invention has the advantages that:
Body and inlet channel coupling are often considered by the hypersonic Waverider aircraft integrated design method of traditional air suction type, namely wish that body and inlet channel are preferably in the shock wave flow field of same intensity, to reduce the transverse flow at inlet mouth place.But the design objective of inlet channel and body is different.Inlet channel needs the object plane by the angle of attack is larger to compress incoming flow, and hypersonic air-flow just can be made to be reduced to the Mach number of combustion chamber needs.Body object plane then needs could produce comparatively high lift-drag ratio under the less angle of attack.If therefore consider that body is identical with the shock strength of inlet channel, be coupled body and Design of Inlet, and body object plane angle must be caused comparatively large, and 1ift-drag ratio is less.The most important meaning of mach front cutting method is can to a certain extent by the design decoupling zero of rider body and inlet channel.Because object plane angle during body generation high lift-drag ratio is generally less than the object plane angle equaling integrated inlet channel precompressed compression face, so inlet channel is after following the trail of generation compressing surface based on mach line, its design objective just considers engine performance parameter and inlet channel startability; The design objective of rider body is then pursue high lift-drag ratio, and the 1ift-drag ratio loss caused for mating punching engine leaves sufficient space.Known according to the above analysis, the shock wave produced due to inlet channel is better than body, therefore in order to high pressure gas after ensureing to follow the trail of inlet channel shock wave that mach line formed do not leak block upper surface, require that body leading edge sweep preferably can be greater than the Angle of Shock Waves in inlet channel one stage of compression face.Before adopting mach front cutting method gained of the present invention, physical efficiency ensures that inlet flow is even, simultaneously equal with inlet channel width at inlet mouth place, avoids the drag due to shock wave increase increasing precursor width and cause, thus improves full machine 1ift-drag ratio.
Accompanying drawing explanation
Fig. 1 is Russian Igla inlet channel schematic diagram;
Fig. 2 is U.S. HSSW aircraft inlet channel schematic diagram;
Fig. 3 is sphenoid and cut sphenoid flow field afterwards along mach front and contrast, and figure conditional is: Ma=5.0, δ=10 °;
Fig. 4 integrated Forebody/Inlet compression system birds-eye view;
Fig. 5 integrated Forebody/Inlet compression system transparent view;
Cloud atlas such as mach line such as grade on the integrated Forebody/Inlet plane of symmetry of Fig. 6;
Cloud atlas such as mach line such as grade on the integrated Forebody/Inlet different cross section of Fig. 7.
In figure:
1. edge, one stage of compression face; 2. edge, two-stage compression face; 3. inlet mouth; 4. edge, three stage compression face; 5. level Four 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. edge in face of one stage of compression.
Detailed description of the invention
Below in conjunction with accompanying drawing embodiment, the present invention is described in detail.
The present invention proposes the mach front cutting method of a kind of hypersonic aircraft Forebody/Inlet integrated design, for two-dimentional wedge shape precursor, according to mach front incision principle, adopt the method for upstream following the trail of mach line to form integrated Forebody/Inlet, ensureing that the drag due to shock wave that inlet flow avoids increase precursor width to bring uniformly simultaneously increases.
The mach front cutting method of hypersonic aircraft Forebody/Inlet provided by the invention integrated design, comprises the following steps:
Step 1: by the ratio of width to height of combustion chamber Demand Design inlet channel cross-sectional.To the two-dimentional inlet channel adopting wedge shape precursor as external compression part, external compression face relies on Two-Dimensional Shock to compress, and the contraction of stream pipe all concentrates on short transverse, therefore combustion chamber can adopt square-section combustion chamber.
Step 2: by the design parameters of inlet channel compression system, determines the incoming flow parameter of inlet channel.
First, according to flight Mach number and engine design Mach number, wave system quantity is determined.Then, organize wave system method by equal strength, determine compressing surface Mach number at different levels, Angle of Shock Waves, air-flow turning angle parameter, the normal component of Shi Ge road oblique shock wave wavefront Mach number is equal, that is:
Ma
1sinβ
1＝Ma
2sinβ
2＝…＝Ma
n-2sinβ
n-2(1)
After shock wave front, Mach number relational expression is:
Wherein, Ma
i, Ma
i+1be respectively Mach number after wavefront, ripple.
The relational expression of Angle of Shock Waves β and air-flow knuckle δ is:
In various above, Ma
ifor Mach number, β
ifor Angle of Shock Waves, γ is specific heats of gases ratios, and δ is air-flow knuckle, i=1,2 ..., n, n are compressing surface progression.
Under the condition that the design parameters (comprising free stream Mach number and inlet mouth Mach number) of inlet channel compression system is given, simultaneous is with above formula (1) ~ (3), solve Mach number after each road shock front, ripple, and compressing surface Mach number at different levels, Angle of Shock Waves, air-flow knuckle etc.
Step 3: according to compressing surface Mach number at different levels, by inlet mouth width and its higher level's compressing surface Mach number, upstream follows the trail of mach line, forms higher level's compressing surface edge.Specifically, if higher level's compressing surface Mach 2 ship Ma
2, then, by inlet mouth marginal point, Mach angle μ=arcsin (1/Ma is made
2) mach line, follow the trail of this mach line and namely form higher level's compressing surface edge.So repeatedly, compressing surface edge at different levels is obtained.Again according to the compressing surface Angle of Shock Waves at different levels of gained in step 2, obtain final external compression/inlet channel compression system.
Fig. 3 gives the shock wave flow field of sphenoid and the contrast in former sphenoid shock wave flow field after mach front cutting, after can finding out excision, disturbance does not propagate into the central area below sphenoid, flowing in central area, without any change, illustrates other regions that can't affect flow field along mach front cutting flow field.
embodiment: design free stream Mach number Ma=5, air intake port Mach number Ma=2.The integrated Forebody/Inlet obtained according to mach front cutting method provided by the invention design as shown in Figure 4 and Figure 5.In this embodiment, inlet mouth 3 is a square-section entrance, and designed compressing surface is level Four compressing surface, and wherein lip puts the third stage and fourth stage compressing surface with lining, and precursor arranges the first order and second stage compressing surface.
Lip with inner compressing surface method of designing is: by inlet mouth 3 lower edge, against carrying out flow path direction, namely upstream level Four compressing surface edge 5 is followed the trail of, obtain fourth stage compressing surface, the position of level Four compressing surface leading edge 6, the shock wave that air-flow should be made to produce after this level Four compressing surface leading edge 6 meets at inlet mouth 3 upper limb; Upstream follow the trail of edge, three stage compression face 4 by level Four compressing surface leading edge 6 again, obtain third stage compressing surface, the shock wave that in face of three stage compression, the position of edge 7 should make air-flow produce after edge 7 in face of this three stage compression meets at inlet mouth 3 upper limb.
The compressing surface method of designing of precursor is: by inlet mouth 3 upper limb, upstream follow the trail of edge, two-stage compression face 2, obtain second stage compressing surface, the shock wave that 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 meets at lip; Upstream follow the trail of edge, one stage of compression face 1 by edge in face of two-stage compression 8 again, obtain first order compressing surface, the shock wave that in face of one stage of compression, the position of edge 9 should make air-flow produce after edge 9 in face of this one stage of compression meets at lip.
As seen from Figure 5, the position sweepforward of gained inlet lip, the compressing surface width upstream following the trail of formation along mach line is greater than inlet mouth width.External compression system both sides are baffled without the need to adding, and also can ensure the uniform incoming flow that inlet channel needs, and both sides can be used as overflow ducts, to improve the starting ability of inlet channel simultaneously.
Adopt Euler equation to carry out numerical modelling to designed inlet channel, calculating total pressure recovery coefficient is 0.8270.Fig. 6 gives the cloud atlas of the mach line such as grade on the integrated Forebody/Inlet plane of symmetry, can find out that the twice shock wave of inlet channel upper wall surface all converges in lower wall surface lip position, the twice shock wave of lower wall surface all converges in inlet mouth, and parameter is even after ripple, show that Theoretical Design coincide with number analog result.Fig. 7 gives second and third grade of compressing surface, and inlet channel leading edge, air intake port cross section Euler equation etc. mach line cloud atlas.Can find out that the precursor of application mach line tracer technique design can provide colory incoming flow for inlet channel.
Claims (3)
1. the mach front cutting method of hypersonic aircraft Forebody/Inlet integrated design, is characterized in that:
Step one, the ratio of width to height by combustion chamber Demand Design inlet channel cross-sectional, to the two-dimentional inlet channel adopting wedge shape precursor as external compression part, combustion chamber adopts square-section combustion chamber;
Step 2, design parameters by inlet channel compression system, determine the incoming flow parameter of inlet channel, and organize wave system method to determine compressing surface Mach number at different levels, Angle of Shock Waves, air-flow turning angle according to equal strength;
The design parameters of described inlet channel compression system comprises 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, form afterbody compressing surface edge; Again by afterbody compressing surface, upstream follow the trail of mach line according to its higher level's compressing surface Mach number, form penultimate stage compressing surface edge; By that analogy, compressing surface edge at different levels is obtained; Again according to the compressing surface Angle of Shock Waves at different levels of gained in step 2, obtain final external compression/inlet channel compression system.
2. the mach front cutting method of hypersonic aircraft Forebody/Inlet according to claim 1 integrated design, is characterized in that: the particular content of step 2 is:
First, according to flight Mach number and engine design Mach number, wave system quantity is determined;
Then, organize wave system method by equal strength, determine compressing surface Mach number at different levels, Angle of Shock Waves, air-flow turning angle parameter, the normal component of Shi Ge road oblique shock wave wavefront Mach number is equal, that is:
Ma
1sinβ
1＝Ma
2sinβ
2＝…＝Ma
n-2sinβ
n-2(1)
After shock wave front, Mach number relational expression is:
The relational expression of Angle of Shock Waves β and air-flow knuckle δ is:
In various above, Ma
i, Ma
i+1be respectively Mach number after the wavefront of i-th grade of compressing surface, ripple, β
ibe the Angle of Shock Waves of i-th grade of compressing surface, γ is specific heats of gases ratios, and δ is air-flow knuckle, i=1,2 ..., n, n are compressing surface progression;
Under the condition that the design parameters of inlet channel compression system is given, simultaneous, with above formula (1) ~ (3), solves Mach number after each road shock front, ripple, and compressing surface Mach number at different levels, Angle of Shock Waves, air-flow knuckle.
3. the mach front cutting method of hypersonic aircraft Forebody/Inlet according to claim 1 integrated design, is characterized in that: the mach line described in step 3 refers to, if higher level's compressing surface Mach 2 ship Ma
2, then, by inlet mouth marginal point, Mach angle μ=arcsin (1/Ma is made
2) mach line.
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| CN105539863A (en) * | 2016-01-29 | 2016-05-04 | 南京航空航天大学 | Integrated aerodynamic layout method for hypersonic aircraft forebody, air inlet duct and supporting plate |
| CN105667811A (en) * | 2016-01-27 | 2016-06-15 | 南京航空航天大学 | Design method for multi-stage coupling integrated structure of front body and air inflow channel of hypersonic aircraft |
| CN105738067A (en) * | 2016-02-01 | 2016-07-06 | 南京航空航天大学 | Rapid determination method for parameters after intersection of two ipsilateral oblique shock waves |
| CN106043738A (en) * | 2016-06-29 | 2016-10-26 | 中国人民解放军国防科学技术大学 | Design method for wide-speed-range waverider aircraft with equal shock wave flow field and variable mach number |
| CN106043737A (en) * | 2016-06-29 | 2016-10-26 | 中国人民解放军国防科学技术大学 | Design method for equal object surface-variable mach number wide-speed-range waverider aircraft |
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| CN105667811A (en) * | 2016-01-27 | 2016-06-15 | 南京航空航天大学 | Design method for multi-stage coupling integrated structure of front body and air inflow channel of hypersonic aircraft |
| CN105539863A (en) * | 2016-01-29 | 2016-05-04 | 南京航空航天大学 | Integrated aerodynamic layout method for hypersonic aircraft forebody, air inlet duct and supporting plate |
| CN105738067A (en) * | 2016-02-01 | 2016-07-06 | 南京航空航天大学 | Rapid determination method for parameters after intersection of two ipsilateral oblique shock waves |
| 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 |
| CN106043738A (en) * | 2016-06-29 | 2016-10-26 | 中国人民解放军国防科学技术大学 | Design method for wide-speed-range waverider aircraft with equal shock wave flow field and variable mach number |
| CN106043737A (en) * | 2016-06-29 | 2016-10-26 | 中国人民解放军国防科学技术大学 | Design method for equal object surface-variable mach number wide-speed-range waverider aircraft |
| 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 |
| 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 |
| CN110162901B (en) * | 2019-05-28 | 2020-03-31 | 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 |
| CN113743033A (en) * | 2021-08-30 | 2021-12-03 | 北京航空航天大学 | Prediction method for height of supersonic jet Mach disk |
| CN113743033B (en) * | 2021-08-30 | 2023-12-12 | 北京航空航天大学 | Prediction method for supersonic jet Mach disk height |
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