CN102996253A - Supersonic air intake duct and wall face determination method of supersonic air intake duct - Google Patents
Supersonic air intake duct and wall face determination method of supersonic air intake duct Download PDFInfo
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Abstract
The invention provides a supersonic air intake duct and a wall face determination method of the supersonic air intake duct. The wall face determination method of the supersonic air intake duct comprises the steps of: determining a front body wall face curve of the supersonic air intake duct according to geometric constraint of structural design of the supersonic air intake duct and by adopting a characteristics method; determining an outer compression wall face curve of the supersonic air intake duct according to an outer contraction ratio and by adopting a second-order continuous curve; determining a lip opening compression curve, a single side wall face curve and an opposite side wall face curve of the supersonic air intake duct according to geometric constraint of structural design of the supersonic air intake duct and by adopting the characteristics method; determining a lower wall face of the supersonic air intake duct according to the front body wall face curve, the outer compression wall face curve and the single side wall face curve; and determining an upper wall face of the supersonic air intake duct according to the lip opening compression curve and the opposite side wall face curve. The supersonic air intake duct determined according to the method is suitable for uneven incoming flow, isentropic compression and shock wave compression of the supersonic air intake duct can be distributed within a certain range through adjusting shock wave shapes, and further total pressure recovery capability can be improved.
Description
Technical field
The present invention relates to the aerodynamics field, determine method in particular to a kind of Supersonic Inlet and wall thereof.
Background technique
Supersonic Inlet is one of air-inlet type ultrasound velocity propulsion system critical component, and the uniformity of the exit flow of intake duct has directly affected the combustion efficiency of motor and then affected the overall performance of aircraft.Existing Supersonic Inlet can't be used for non-homogeneous inlet flow conditions, and the pitot loss of many shock wave compressions intake duct is larger, and compression efficiency is lower, and flow coefficient is larger with change in angle of attack, is unfavorable for the wide range work of aircraft.Interior pressure channel design is stronger with precursor design coupling, is difficult to dock different firing chambers.Therefore under the same regime of flight of examination, during different chamber performance, need a plurality of intake ducts of design to connect with it, both expended a large amount of manpower and materials, be difficult to again shorten experimental period.
Existing Supersonic Inlet Design method has a variety of, and wherein the most typical is the Design of Inlet method of a kind of many multishocks of monograph " aircraft internal aerodynamics " discussion, and its step is as follows:
(1) according to the Choice and design requirement of design point, determines precursor shock wave number;
(2) according to the requirement of total pressure recovery coefficient, determine the angle of precursor compressing surfaces at different levels;
(3) according to the requirement of interior compression profile, determine labial angle in the outer cover;
(4) according to the flow coefficient requirement, determine that compressing surface is with respect to the position of lip;
(5) carry out the three-dimensional structure design.
The intake duct of existing many multishocks can't be used for non-homogeneous inlet flow conditions, and the pitot loss of this intake duct is larger, at the bottom of the compression efficiency, flow coefficient is larger with change in angle of attack, is unfavorable for the work of the wide range of aircraft.And existing many shock waves Design of Inlet and precursor design coupling are stronger, be difficult to dock different firing chambers, therefore, under the same regime of flight of examination during different chamber performance, need a plurality of intake ducts of design to connect with it, not only be difficult to shorten experimental period, and expended a large amount of manpower and materials.
Summary of the invention
The present invention aim to provide a kind ofly be suitable for non-homogeneous incoming flow, Supersonic Inlet and wall thereof that pitot loss is little are determined method.
To achieve these goals, according to an aspect of the present invention, provide a kind of wall of Supersonic Inlet to determine method, having comprised: according to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the precursor wall curve of Supersonic Inlet; According to outer contraction ratio, utilize the Second Order Continuous curve, determine the external compression wall curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the lip compression curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the monolateral wall curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the opposite side wall curve corresponding to monolateral wall curve; Precursor wall curve, external compression wall curve and monolateral wall curve are determined the lower wall surface of Supersonic Inlet; Lip compression curve and opposite side wall curve are determined the upper wall surface of Supersonic Inlet.
Further, according to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine that the step of the precursor wall curve of Supersonic Inlet comprises: according to forebody length with catch requirement for height, set precursor shock wave curve.
Further, set after the precursor shock wave curve, geometric constraint according to the Supersonic Inlet structural design, adopt method of characteristics, the step of determining the precursor wall curve of Supersonic Inlet also comprises: according to the shape of flying condition and precursor shock wave curve, determine flow field behind precursor wall curve and the ripple.
Further, determine after the precursor wall curve, according to outer contraction ratio, utilize the Second Order Continuous curve, the step of determining the external compression wall curve of Supersonic Inlet comprises: according to outer contraction ratio, determine the height of Supersonic Inlet entrance, take precursor shock wave curve away from the second end points of entrance as the center of circle, justify take the height of Supersonic Inlet entrance as radius, utilize the end points away from precursor shock wave curve of Second Order Continuous curve the past body wall surface curve to begin to determine external compression wall curve, and make this external compression wall curve tangent with circle.
Further, geometric constraint according to the Supersonic Inlet structural design, adopt method of characteristics, the step of determining the lip compression curve of Supersonic Inlet comprises: the characteristic line and the external compression curve that form according to the end points away from precursor wall curve away from the end points of precursor shock wave curve and precursor shock wave curve of precursor wall curve, adopt method of characteristics, determine the external compression zone, wherein, the boundary point in this external compression zone is respectively precursor shock wave curve away from the end points of precursor wall curve and the two-end-point of external compression wall curve.
Further, determine after the external compression zone, geometric constraint according to the Supersonic Inlet structural design, adopt method of characteristics, the step of determining the lip compression curve of Supersonic Inlet also comprises: according to the designing requirement of inside and outside resistance, determine the lip compression curve, adopt method of characteristics to determine the lip shock flow field according to this lip compression curve, and the intersection point of definite lip shock flow field and external compression wall curve, wherein, the lip compression curve end points away from precursor wall curve that end points is the shock wave curve.
Further, determine after the lip shock curve, geometric constraint according to the Supersonic Inlet structural design, adopt method of characteristics, the monolateral wall curve of determining Supersonic Inlet comprises: according to interior compression geometric constraint, adopt method of characteristics, determine monolateral wall curve, wherein, monolateral wall curve intersection point that end points is lip shock flow field and external compression wall curve.
Further, the step of determining the monolateral wall curve of Supersonic Inlet according to the geometric constraint of Supersonic Inlet structural design also comprises: when determining monolateral wall curve, the Mach Number Distribution of this single wall surface curve is set, so that the single wall surface curve overlaps with local flow direction angle with the angle of contingence at the end points place of the intersection point of external compression wall curve in the lip shock flow field.
Further, utilize method of characteristics to determine to comprise corresponding to the step of the opposite side wall curve of monolateral wall curve: according to the Mach Number Distribution of monolateral wall curve, utilize method of characteristics, find the solution the opposite side wall curve of monolateral wall curve, an end of this opposite side wall curve is the end points away from precursor wall curve of shock wave curve.
Further, method of characteristics comprises estimates step and the step of correction, proofreaies and correct the step and proofreaies and correct according to estimating the result in step.
According to a further aspect in the invention, provide a kind of Supersonic Inlet, the wall of Supersonic Inlet determines that by the wall of above-mentioned Supersonic Inlet method is definite.
In accordance with a further aspect of the present invention, a kind of Supersonic Inlet is provided, Supersonic Inlet comprises upper wall surface, lower wall surface and be connected to upper wall surface and lower wall surface between two side wall surfaces, lower wall surface is by precursor wall curve, external compression wall curve, and monolateral wall curve is determined, upper wall surface is determined with the opposite side wall curve corresponding with monolateral wall curve by the lip compression curve, wherein, precursor wall curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design, external compression wall curve utilizes the Second Order Continuous curve to determine by outer contraction ratio, the lip compression curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design, and monolateral wall curve opposite side wall curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design.
Use technological scheme of the present invention, the wall of Supersonic Inlet determines that method comprises: according to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the precursor wall curve of Supersonic Inlet wall curve; According to outer contraction ratio, utilize the Second Order Continuous curve, determine the external compression wall curve of Supersonic Inlet wall curve; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the lip compression curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the monolateral wall curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the opposite side wall curve corresponding to monolateral wall curve.The Supersonic Inlet that the method according to this invention is determined is applicable to non-homogeneous incoming flow, and its isentropic Compression and shock wave compression can distribute within the specific limits by adjusting profile of shock wave, thereby improves total pressure recovery.Simultaneously, this Supersonic Inlet can adjust to dock different motors flexibly, can also guarantee flow field quality simultaneously.
Description of drawings
The accompanying drawing that consists of a part of the present invention is used to provide a further understanding of the present invention, and illustrative examples of the present invention and explanation thereof are used for explaining the present invention, do not consist of improper restriction of the present invention.In the accompanying drawings:
Fig. 1 shows according to Supersonic Inlet of the present invention and wall thereof and determines the precursor shock wave curve that method forms;
Fig. 2 shows according to Supersonic Inlet of the present invention and wall thereof and determines flow field behind precursor wall curve that method forms and the ripple;
Fig. 3 shows according to Supersonic Inlet of the present invention and wall thereof and determines the external compression curve that method forms;
Fig. 4 shows according to Supersonic Inlet of the present invention and wall thereof and determines the external compression zone that method forms;
Fig. 5 shows according to Supersonic Inlet of the present invention and wall thereof and determines lip compression curve and the lip shock flow field that method forms;
Fig. 6 shows according to Supersonic Inlet of the present invention and wall thereof and determines the monolateral wall curve that method forms;
Fig. 7 show according to Supersonic Inlet of the present invention and wall thereof determine that method forms in the relative opposite side wall curve of monolateral wall curve;
Fig. 8 shows the schematic diagram of determining other Supersonic Inlets that method forms according to Supersonic Inlet of the present invention and wall thereof; And
Fig. 9 shows the solution procedure of determining the characteristic line equation of method according to Supersonic Inlet of the present invention and wall thereof.
Embodiment
Hereinafter also describe in conjunction with the embodiments the present invention in detail with reference to accompanying drawing.Need to prove, in the situation that do not conflict, embodiment and the feature among the embodiment among the application can make up mutually.
Among the present invention, the Mach number of ultrasound velocity section is greater than 1.2, and the Mach number of subsonic velocity section is less than 0.8, and transonic speed the Mach number of section is between 0.8 to 1.2.
According to embodiments of the invention, the wall of Supersonic Inlet obtains by the following method.
As shown in Figure 1, at first according to forebody length with catch requirement for height, set precursor shock wave curve 1-2, this shock wave curve 1-2 is bendable directly, to reach specific compression purpose, wherein, forebody length refers to that the intake duct leading edge point is apart from lip x direction distance under air path axis system.Catch and refer to that highly the intake duct leading edge point is apart from lip y direction distance under air path axis system.As shown in Figure 2, set after the precursor shock wave curve 1-2, according to the shape of flying condition and precursor shock wave curve 1-2, utilize method of characteristics, determine flow field 1-2-3 behind precursor wall curve 1-3 and the ripple, wherein flying condition be pressure, height and the Mach number of locality.
As shown in Figure 3, determine behind precursor wall curve 1-3 and the ripple after the 1-2-3 of flow field, according to outer contraction ratio, namely catch height and the ratio of inner flow passage entrance height, determine the inlet mouth height H, take precursor shock wave curve 1-2 away from the end points 2 of entrance as the center of circle, H is that radius is determined circle C, adopt the Second Order Continuous curve, the end points 3 away from precursor shock wave curve 1-2 of body wall surface curve 1-3 begins to determine external compression wall curve 3-4 in the past, wherein, 4 among the external compression curve 3-4 is free point, its distance with point 3 needs only and can guarantee that lip shock flow field 2-5-6 and the external compression curve 3-4 that will find the solution intersect, and this external compression wall curve 3-4 is tangent with circle C.
As shown in Figure 4, determine after the external compression curve 3-4, the characteristic line 2-3 and the external compression curve 3-4 that form according to the end points 2 away from precursor wall curve 1-3 away from the end points 3 of precursor shock wave curve 1-2 and precursor shock wave curve 1-2 of precursor wall curve 1-3, adopt method of characteristics, determine external compression zone 2-3-4, wherein, the boundary point of this external compression zone 2-3-4 is respectively precursor shock wave curve 1-2 away from the end points 2 of precursor wall curve 1-3 and the two-end-point of external compression wall curve 3-4, i.e. end points 3 and end points 4.
Determine as shown in Figure 5 after the described external compression zone 2-3-4, designing requirement according to inside and outside resistance, determine lip compression curve 2-6, and then employing method of characteristics, determine lip shock flow field 2-5-6, wherein, lip shock flow field 2-5-6 and external compression wall curve 3-4 meet at a little 5, and the end points of lip compression curve 2-6 is that precursor shock wave curve 1-2 is away from the end points 2 of precursor wall curve 1-3.Here said interior extrernal resistance designing requirement refers to the design restriction of given inside and outside resistance, can be a numerical value of determining according to the Aircraft Conceptual Design requirement.
As shown in Figure 6, determine after lip compression curve 2-6 and the lip shock flow field 2-5-6, according to interior compression geometric constraint, the length of compression, height etc. namely, determine monolateral wall curve 5-7, wherein, the end points of monolateral wall curve 5-7 is through the intersection point 5 of lip shock flow field 2-5-6 and external compression wall curve 3-4.When determining described monolateral wall curve 5-7, the Mach Number Distribution of this single wall surface curve 5-7 is set, so that single wall surface curve 5-7 overlaps with local flow direction angle in the angle of contingence through 5 place, can guarantees the shock wave areflexia like this.Wherein, the angle of contingence is the tangent line of curve and the angle of x axle; Flow angle is the angle of flow direction and x axle.
As shown in Figure 7, determine according to the Mach Number Distribution of monolateral wall curve 5-7, to utilize method of characteristics after the monolateral wall curve 5-7, find the solution the opposite side wall curve 6-8 relative with monolateral wall curve 5-7, the end points 6 of this opposite side wall curve links to each other with the point 6 of lip compression curve 2-6.
As shown in Figure 8, compression geometric constraint in changing, repeat to determine the process of monolateral wall curve 5-7 and the opposite side wall curve 6-8 relative with monolateral wall curve 5-7, thereby obtain new monolateral wall curve 5-7 ' and new opposite side wall curve 6-8 ', to dock different motors, finally determine the wall of whole Supersonic Inlet.
The process of wherein utilizing method of characteristics that the wall curve is found the solution is as follows:
Suppose two point (x on the known curve
1, r
1, M
1, θ
1), (x
2, r
2, M
2, θ
2), need to find the solution thirdly (x
3, r
3, M
3, θ
3) time, can utilize process shown in Figure 9 to find the solution.
In solution procedure, at first according to estimating the step to thirdly finding the solution, then the value of finding the solution is proofreaied and correct coordinate thirdly, Mach number and flow direction angle after obtaining to proofread and correct.
Estimate to go on foot and comprise:
Find the solution first (x
3, r
3),
μ
1＝sin
-1(1/M
1)
μ
2＝sin
-1(1/M
2)
h
1=tan[θ
1+μ
1]
h
2=tan[θ
2-μ
2]
Have according to difference equation:
r
3-r
1=h
1(x
3-x
1)
r
3-r
2=h
2(x
3-x
2)
Two formulas are subtracted each other and can be got:
r
1-r
2={h
2-h
1}x
3+x
1h
1-x
2h
2
Try to achieve coordinate thirdly
The below finds the solution the compatibility relation formula:
Order:
Then have:
g
1(M
3-M
1)-(θ
3-θ
1)-f
1=0
g
2(M
3-M
2)+(θ
3-θ
2)-f
2=0
Thereby obtain thirdly Mach number and the flow direction angle at place, position:
θ
3=g
1(M
3-M
1)+θ
1-f
1
μ
3＝sin
-1(1/M
3)
In the above-mentioned formula, M
1Be the Mach number at place, first position, μ
1Be the Mach angle at place, first position, θ
1Be the flow direction angle at place, first position, x
1Be the abscissa at place, first position, r
1Be the y coordinate at place, first position, γ is the specific heat at constant pressure of gas and the ratio of specific heat of specific heat at constant volume, and M is local Mach number and M〉1, δ is the pattern of flow parameter, for two-dimensional flow δ=0, Three-dimensional Axisymmetric δ=1 of flowing, r ≠ 0.
M
2Be the Mach number at place, second point position, μ
2Be the Mach angle at place, second point position, θ
2Be the flow direction angle at place, second point position, x
2Be the abscissa at place, second point position, r
2Y coordinate for place, second point position.
M
3Be the Mach number at place, position thirdly, μ
3Be the Mach angle at place, position thirdly, θ
3Be the flow direction angle at place, position thirdly, x
3Be the abscissa at place, position thirdly, r
3Y coordinate for place, position thirdly.
After estimating coordinate, Mach number and the flow direction angle that solves place, position thirdly in the step, the coefficient of equation or parameter averaged repeat to estimate the computational process in step, thirdly Mach number and flow direction angle are proofreaied and correct.This parameter or coefficient mean value can be found the solution by Mach number thirdly and the flow direction angle of trying to achieve, order
M wherein
1' be first the Mach number mean value after proofreading and correct, M
2' for the Mach number mean value of second point after proofreading and correct, then with M
1' and M
2' the value substitution estimate and proceed in the step to find the solution, until the correction of a final proof step thirdly Mach number of trying to achieve with estimate the thirdly Mach number M that tries to achieve in the step
3Location of equal, the final Mach number after the Mach 2 ship at the place, thirdly present position of this moment is proofreaied and correct.In like manner, thirdly the flow direction angle at place, position also can obtain final flow direction angle by proofreading and correct the step.
According to embodiments of the invention, a kind of Supersonic Inlet is provided, Supersonic Inlet comprises upper wall surface, lower wall surface and be connected to upper wall surface and lower wall surface between two side wall surfaces, lower wall surface is by precursor wall curve, external compression wall curve, and monolateral wall curve is determined, upper wall surface is determined with the opposite side wall curve corresponding with monolateral wall curve by the lip compression curve, wherein, precursor wall curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design, external compression wall curve utilizes the Second Order Continuous curve to determine by outer contraction ratio, the lip compression curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design, and monolateral wall curve opposite side wall curve adopts method of characteristics to determine by the geometric constraint of Supersonic Inlet structural design.
From above description, can find out, the above embodiments of the present invention have realized following technique effect: the wall of Supersonic Inlet determines that method comprises: according to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the precursor wall curve of Supersonic Inlet wall curve; According to outer contraction ratio, utilize the Second Order Continuous curve, determine the external compression wall curve of Supersonic Inlet wall curve; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the lip compression curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the monolateral wall curve of Supersonic Inlet; According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the opposite side wall curve corresponding to monolateral wall curve.The Supersonic Inlet that the method according to this invention is determined is applicable to non-homogeneous incoming flow, and its isentropic Compression and shock wave compression can distribute within the specific limits by adjusting profile of shock wave, thereby improves total pressure recovery.Simultaneously, this Supersonic Inlet can adjust to dock different motors flexibly, can also guarantee flow field quality simultaneously.
The above is the preferred embodiments of the present invention only, is not limited to the present invention, and for a person skilled in the art, the present invention can have various modifications and variations.Within the spirit and principles in the present invention all, any modification of doing, be equal to replacement, improvement etc., all should be included within protection scope of the present invention.
Claims (12)
1. the wall of a Supersonic Inlet is determined method, it is characterized in that, comprising:
According to the geometric constraint of Supersonic Inlet structural design, adopt method of characteristics, determine the precursor wall curve of described Supersonic Inlet;
According to outer contraction ratio, utilize the Second Order Continuous curve, determine the external compression wall curve of described Supersonic Inlet;
According to the geometric constraint of described Supersonic Inlet structural design, adopt described method of characteristics, determine the lip compression curve of Supersonic Inlet;
According to the geometric constraint of described Supersonic Inlet structural design, adopt described method of characteristics, determine the monolateral wall curve of described Supersonic Inlet;
According to the geometric constraint of described Supersonic Inlet structural design, adopt described method of characteristics, determine the opposite side wall curve corresponding to described monolateral wall curve;
Described precursor wall curve, external compression wall curve and monolateral wall curve are determined the lower wall surface of described Supersonic Inlet; Described lip compression curve and described opposite side wall curve are determined the upper wall surface of described Supersonic Inlet.
2. the wall of Supersonic Inlet according to claim 1 is determined method, it is characterized in that, described geometric constraint according to described Supersonic Inlet structural design adopts described method of characteristics, determines that the step of the precursor wall curve of described Supersonic Inlet comprises:
According to forebody length and catch requirement for height, set described precursor shock wave curve.
3. the wall of Supersonic Inlet according to claim 2 is determined method, it is characterized in that, set after the described precursor shock wave curve, described geometric constraint according to described Supersonic Inlet structural design, adopt described method of characteristics, determine that the step of the precursor wall curve of described Supersonic Inlet also comprises:
According to the shape of flying condition and described precursor shock wave curve, determine flow field behind described precursor wall curve and the ripple.
4. the wall of Supersonic Inlet according to claim 3 is determined method, it is characterized in that, determine after the described precursor wall curve, described according to outer contraction ratio, utilize described Second Order Continuous curve, determine that the step of the external compression wall curve of described Supersonic Inlet comprises:
According to described outer contraction ratio, determine the height of described Supersonic Inlet entrance, take described precursor shock wave curve away from the second end points of entrance as the center of circle, justify take the height of described Supersonic Inlet entrance as radius, utilize described Second Order Continuous curve to begin to determine described external compression wall curve from the end points away from described precursor shock wave curve of described precursor wall curve, and make this external compression wall curve and described circle tangent.
5. the wall of Supersonic Inlet according to claim 4 is determined method, it is characterized in that, described geometric constraint according to described Supersonic Inlet structural design adopts described method of characteristics, determines that the step of the lip compression curve of described Supersonic Inlet comprises:
The characteristic line and the described external compression curve that form according to the end points away from described precursor wall curve away from the end points of described precursor shock wave curve and precursor shock wave curve of described precursor wall curve, adopt method of characteristics, determine the external compression zone, wherein, the boundary point in this external compression zone is respectively described precursor shock wave curve away from the end points of described precursor wall curve and the two-end-point of described external compression wall curve.
6. the wall of Supersonic Inlet according to claim 5 is determined method, it is characterized in that, determine after the described external compression zone, described geometric constraint according to described Supersonic Inlet structural design, adopt described method of characteristics, determine that the step of the lip compression curve of described Supersonic Inlet also comprises:
Designing requirement according to inside and outside resistance, determine described lip compression curve, adopt described method of characteristics to determine the lip shock flow field according to this lip compression curve, and the intersection point of definite described lip shock flow field and described external compression wall curve, wherein, described lip compression curve end points away from described precursor wall curve that end points is described shock wave curve.
7. the wall of Supersonic Inlet according to claim 6 is determined method, it is characterized in that, determine after the described lip shock curve, described geometric constraint according to described Supersonic Inlet structural design, adopt described method of characteristics, determine that the monolateral wall curve of described Supersonic Inlet comprises:
According to interior compression geometric constraint, adopt described method of characteristics, determine described monolateral wall curve, wherein, an end points of described monolateral wall curve is the intersection point of described lip shock flow field and described external compression wall curve.
8. the wall of Supersonic Inlet according to claim 7 is determined method, it is characterized in that, described geometric constraint according to the Supersonic Inlet structural design determines that the step of the monolateral wall curve of Supersonic Inlet also comprises:
When determining described monolateral wall curve, the Mach Number Distribution of this single wall surface curve is set, so that described single wall surface curve overlaps with local flow direction angle with the angle of contingence at the end points place of the intersection point of described external compression wall curve in described lip shock flow field.
9. the wall of Supersonic Inlet according to claim 8 is determined method, it is characterized in that, describedly utilizes described method of characteristics to determine to comprise corresponding to the step of the opposite side wall curve of monolateral wall curve:
According to the Mach Number Distribution of described monolateral wall curve, utilize described method of characteristics, find the solution the opposite side wall curve of described monolateral wall curve, an end of this opposite side wall curve is the end points away from described precursor wall curve of described shock wave curve.
According to claim 1 in 9 the wall of each described Supersonic Inlet determine method, it is characterized in that, described method of characteristics comprises estimates step and the step of correction, the described correction step proofreaies and correct according to the described result who estimates the step.
11. a Supersonic Inlet is characterized in that, the wall of described Supersonic Inlet determines that by the wall of each described Supersonic Inlet in the claim 1 to 10 method is definite.
12. Supersonic Inlet, it is characterized in that, described Supersonic Inlet comprises upper wall surface, lower wall surface and be connected to described upper wall surface and described lower wall surface between two side wall surfaces, described lower wall surface is by precursor wall curve, external compression wall curve, and monolateral wall curve is determined, described upper wall surface is determined with the opposite side wall curve corresponding with described monolateral wall curve by the lip compression curve, wherein, described precursor wall curve adopts method of characteristics to determine by the geometric constraint of described Supersonic Inlet structural design, described external compression wall curve utilizes the Second Order Continuous curve to determine by outer contraction ratio, described lip compression curve adopts method of characteristics to determine by the geometric constraint of described Supersonic Inlet structural design, and described monolateral wall curve opposite side wall curve adopts described method of characteristics to determine by the geometric constraint of described Supersonic Inlet structural design.
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| CN112944396A (en) * | 2021-05-13 | 2021-06-11 | 中国人民解放军国防科技大学 | Method for measuring mixing efficiency of wall-injected gaseous fuel and supersonic incoming flow |
| CN113800001A (en) * | 2021-09-30 | 2021-12-17 | 西安航天动力研究所 | Design method of internal shrinkage hypersonic inlet channel integrated with forebody |
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| CN103306820A (en) * | 2013-06-28 | 2013-09-18 | 中国人民解放军国防科学技术大学 | Supersonic air feeding channel and three-dimensional determining method for wall surface of supersonic air feeding channel |
| CN103306820B (en) * | 2013-06-28 | 2015-09-16 | 中国人民解放军国防科学技术大学 | The three-dimensional defining method of Supersonic Inlet and wall thereof |
| CN104863716A (en) * | 2015-04-24 | 2015-08-26 | 南京航空航天大学 | Design method for control measure of oblique shock wave/boundary layer interaction in air inlet on basis of binary bulge |
| CN104989549A (en) * | 2015-05-27 | 2015-10-21 | 中国人民解放军装备学院 | Method for raising air inflow yield of ramjet through laser energy injection |
| CN104989549B (en) * | 2015-05-27 | 2020-03-06 | 中国人民解放军战略支援部队航天工程大学 | Method for improving air intake capture of ramjet by injecting laser energy |
| CN112944396A (en) * | 2021-05-13 | 2021-06-11 | 中国人民解放军国防科技大学 | Method for measuring mixing efficiency of wall-injected gaseous fuel and supersonic incoming flow |
| CN112944396B (en) * | 2021-05-13 | 2021-07-09 | 中国人民解放军国防科技大学 | Method for measuring mixing efficiency of wall-injected gaseous fuel and supersonic incoming flow |
| CN113800001A (en) * | 2021-09-30 | 2021-12-17 | 西安航天动力研究所 | Design method of internal shrinkage hypersonic inlet channel integrated with forebody |
| CN113800001B (en) * | 2021-09-30 | 2024-02-27 | 西安航天动力研究所 | Design method of inner-shrinkage hypersonic air inlet channel integrated with precursor |
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