CN108182319A - A kind of supersonic speed integration Nozzle Design method - Google Patents
A kind of supersonic speed integration Nozzle Design method Download PDFInfo
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Abstract
A kind of supersonic speed integration Nozzle Design method, is related to wind tunnel test field；Include the following steps：Step (1) establishes jet pipe model, including contraction section, expansion arc, test section and boundary layer；Step (2) acquires TG sections of curves according to the characteristic strips equation of TG sections of curves；AD sections of curves are acquired according to the characteristic strips equation of AD sections of curves, obtain the complete curve of expansion arc outer wall；Step (3), the displacement thickness for calculating boundary layer；Obtain boundary layer curve；Step (4), the complete curve for obtaining contraction section；Step (5), the complete curve for obtaining test section；The invention enables test area increase, and test section Mach number root-mean-square-deviation, the axial parameters such as Mach number gradient and flow-deviation angle meet the advanced index of national military standard.
Description
Technical field
The present invention relates to a kind of wind tunnel test field, particularly a kind of supersonic speed integration Nozzle Design method.
Background technology
At present in supersonic wind tunnel experiment, since nozzle exit Mach angle is smaller, cause effectively to test homogeneity range face
Product is smaller, in order to avoid test model is exposed to outside experiment homogeneity range, can only use compared with small test model.Smaller test model
The test data and practical flight device model obtained in wind-tunnel obtains test data, and there are certain deviations, influence experiment essence
Degree.Meanwhile smaller experiment homogeneity range, limited model continuous transformation attitude angle improve wind tunnel operation cost.
Invention content
It is an object of the invention to overcome the above-mentioned deficiency of the prior art, a kind of supersonic speed integration Nozzle Design side is provided
Method so that test area increases, and test section Mach number root-mean-square-deviation, the axial parameters such as Mach number gradient and flow-deviation angle
Meet the advanced index of national military standard.
The above-mentioned purpose of the present invention is achieved by following technical solution：
A kind of supersonic speed integration Nozzle Design method, includes the following steps：
Step (1) establishes jet pipe model；Including contraction section, expansion arc, test section and boundary layer；Wherein, contraction section, swollen
Swollen section joins end to end successively in an axial direction with test section；Boundary layer is coated on the outer wall of expansion arc and test section；Contraction section and expansion
The junction of section is venturi；The position that venturi corresponds to jet pipe outer wall is T points；The junction of expansion arc and test section is corresponded to outside jet pipe
The position of wall is D points；Contraction section is arrival end axially away from one end of expansion arc；Expansion arc is axially away from one end of contraction section
The port of export；
Step (2) presets maximum swelling angle θ, the Mach number M of expansion arc and the port of export diameter of expansion arc；It chooses
The point is defined as G points by maximum swelling angle θ corresponding points in expansion arc；Using G points as starting point, the direction along point test section exists
The outer wall of expansion arc takes point A；GA sections are conic line segment；Establish coordinate system oxy；AD sections of songs are established according to coordinate system oxy respectively
The characteristic strips equation of characteristic strips equation and TG section curve of line；TG sections of curves are acquired according to the characteristic strips equation of TG sections of curves；Root
AD sections of curves are acquired according to the characteristic strips equation of AD sections of curves；
The G points of TG sections of curves are connected with the A point straight lines of AD sections of curves；Obtain the complete curve of expansion arc outer wall；
Step (3) establishes toll bar momentum integral equation calculating momentum thickness of boundary layer δ '；And according to momentum thickness δ '
Calculate the displacement thickness δ in boundary layer；As boundary layer curve；
The characteristic strips equation of step (4), TG section curves in step (2), obtains the slope in T points downstream；Another T points
The slope in downstream is identical with the slope of T points upstream；Obtain the complete curve of contraction section；
Step (5) presets and owes expansion modifying factor λ；The outer wall curve of test section is straightway；According to step
(2) characteristic strips equation of AD sections of curves in obtains the slope of D points upstream；The slope of test section is multiplied by for the slope of D points upstream
Owe expansion modifying factor λ；Obtain the complete curve of test section.
In a kind of above-mentioned supersonic speed integration Nozzle Design method, the step (two), the foundation of coordinate system oxy
Method is：
Using T points as coordinate origin, positive direction of the x-axis is side straight up to be axially directed to outlet extreme direction, y-axis square
To.
In a kind of above-mentioned supersonic speed integration Nozzle Design method, the step (two), the feature of AD sections of curves
Line equation is：
The characteristic strips equation of TG sections of curves is：
In formula, x is the abscissa of coordinate system oxy；
Y is the ordinate of coordinate system oxy；
Specific heats of gases ratio in γ jet pipes.
In a kind of above-mentioned supersonic speed integration Nozzle Design method, the step (three), accumulated by toll bar momentum
The method of point equation calculation momentum thickness of boundary layer δ ' is：
Wherein, H is boundary-layer form factor；
CfFor friction coefficient.
In a kind of above-mentioned supersonic speed integration Nozzle Design method, the step (three), the calculating of displacement thickness δ
Method is：
δ=H δ ' (4).
The present invention has the following advantages that compared with prior art：
(1) present invention glues molded line using the nothing of step 1 characteristic line method design, and without viscous molded line slope rate continuity, air-flow is spraying
Flowing will not generate wave system during Bottomhole pressure；
(2) present invention uses the boundary layer correction method of step 2, meets the flowing law of air-flow, the flow field of nozzle exit
It is best in quality；
(3) present invention uses the connection mode of step 3 so that jet pipe contraction section and expansion arc slope rate continuity, air-flow flow through
During venturi T points, flowing will not generate separation；
(4) present invention uses the optimization method of step 4 and step 5 so that nozzle exit DQ does not expand wave system significantly
And shock wave so that test area increases to CDSN by CDF.Test section Mach number root-mean-square-deviation, axial Mach number gradient are gentle
The parameters such as stream deflection angle meet the advanced index of national military standard.
Description of the drawings
Fig. 1 is supersonic speed integration jet pipe schematic diagram of the present invention；
Fig. 2 is supersonic speed integration Nozzle Design flow diagram of the present invention.
Specific embodiment
The present invention is described in further detail in the following with reference to the drawings and specific embodiments：
Supersonic speed integration Nozzle Design flow diagram is illustrated in figure 2, as seen from the figure, a kind of supersonic speed integration spray
Pipe design method, includes the following steps：
Step (1) establishes jet pipe model；Including contraction section 1, expansion arc 2, test section 3 and boundary layer 4；Wherein, it shrinks
Section 1, expansion arc 2 and test section 3 join end to end successively in an axial direction；Boundary layer 4 is coated on the outer wall of expansion arc 2 and test section 3；It receives
The junction of contracting section 1 and expansion arc 2 is venturi；It is as shown in Figure 1 supersonic speed integration jet pipe schematic diagram, as seen from the figure, venturi
The position of corresponding jet pipe outer wall is T points；The position that the junction of expansion arc 2 and test section 3 corresponds to jet pipe outer wall is D points；It shrinks
Section 1 is arrival end axially away from one end of expansion arc 2；Expansion arc 2 is the port of export axially away from one end of contraction section 1；
Step (2) presets maximum swelling angle θ, the Mach number M of expansion arc 2 and the port of export diameter of expansion arc 2；Choosing
Maximum swelling angle θ corresponding points in expansion arc 2 are taken, which is defined as G points；Using G points as starting point, along the side of point test section 3
Point A is taken to the outer wall in expansion arc 2；GA sections are conic line segment；Establish coordinate system oxy；Wherein, the foundation side of coordinate system oxy
Method is：
Using T points as coordinate origin, positive direction of the x-axis is side straight up to be axially directed to outlet extreme direction, y-axis square
To.
Establish the characteristic strips equation of AD sections of curves and the characteristic strips equation of TG sections of curves respectively according to coordinate system oxy；According to
The characteristic strips equation of TG sections of curves acquires TG sections of curves；AD sections of curves are acquired according to the characteristic strips equation of AD sections of curves；
The characteristic strips equation of AD sections of curves is：
The characteristic strips equation of TG sections of curves is：
In formula, x is the abscissa of coordinate system oxy；
Y is the ordinate of coordinate system oxy；
Specific heats of gases ratio in γ jet pipes.
The G points of TG sections of curves are connected with the A point straight lines of AD sections of curves；Obtain the complete curve of 2 outer wall of expansion arc；
Step (3) establishes the momentum thickness δ ' that toll bar momentum integral equation calculates boundary layer 4；And according to momentum thickness δ '
Calculate the displacement thickness δ in boundary layer 4；As 4 curve of boundary layer；
The method of momentum thickness δ ' that boundary layer (4) are calculated by toll bar momentum integral equation is：
Wherein, H is boundary-layer form factor；
CfFor friction coefficient.
The computational methods of displacement thickness δ are：
δ=H δ ' (4).
The characteristic strips equation of step (4), TG section curves in step (2), obtains the slope in T points downstream；Another T points
The slope in downstream is identical with the slope of T points upstream；Obtain the complete curve of contraction section 1；
Step (5) presets and owes expansion modifying factor λ；The outer wall curve of test section 3 is straightway；According to step
(2) characteristic strips equation of AD sections of curves in obtains the slope of D points upstream；The slope of test section 3 is multiplied by for the slope of D points upstream
Owe expansion modifying factor λ；Obtain the complete curve of test section 3.Modifying factor λ is optimized so that nozzle exit D and experiment
Section DS completes the Profile Design of supersonic speed integration jet pipe without apparent dilatational wave and shock wave.
Point C, CD is taken to be equal to 2 port of export radius of expansion arc divided by sin β at nozzle axis；β=arcsin (1/M)；F and C
Point exports radial symmetric along expansion arc 2；The tail end of test section 3 radially takes two points of S, N respectively；The present invention by test area by
CDF increases to CDSN.Test section Mach number root-mean-square-deviation, the axial parameters such as Mach number gradient and flow-deviation angle meet army of state
Mark advanced index.
The content not being described in detail in description of the invention belongs to the known technology of those skilled in the art.
Claims (5)
- A kind of 1. supersonic speed integration Nozzle Design method, it is characterised in that：Include the following steps：Step (1) establishes jet pipe model；Including contraction section (1), expansion arc (2), test section (3) and boundary layer (4)；Wherein, Contraction section (1), expansion arc (2) and test section (3) join end to end successively in an axial direction；Boundary layer (4) is coated on expansion arc (2) and examination Test the outer wall of section (3)；The junction of contraction section (1) and expansion arc (2) is venturi；The position that venturi corresponds to jet pipe outer wall is T points； The position that expansion arc (2) corresponds to jet pipe outer wall with the junction of test section (3) is D points；Contraction section (1) is axially away from expansion arc (2) one end is arrival end；Expansion arc (2) is the port of export axially away from one end of contraction section (1)；The port of export diameter of step (2), the maximum swelling angle θ, the Mach number M that preset expansion arc (2) and expansion arc (2)；Choosing Maximum swelling angle θ corresponding points in expansion arc (2) are taken, which is defined as G points；Using G points as starting point, along point test section (3) Direction take point A in the outer wall of expansion arc (2)；GA sections are conic line segment；Establish coordinate system oxy；Distinguished according to coordinate system oxy Establish the characteristic strips equation of AD sections of curves and the characteristic strips equation of TG sections of curves；TG is acquired according to the characteristic strips equation of TG sections of curves Section curve；AD sections of curves are acquired according to the characteristic strips equation of AD sections of curves；The G points of TG sections of curves are connected with the A point straight lines of AD sections of curves；Obtain the complete curve of expansion arc (2) outer wall；Step (3) establishes the momentum thickness δ ' that toll bar momentum integral equation calculates boundary layer (4)；And according to momentum thickness δ ' meters Calculate the displacement thickness δ of boundary layer (4)；As boundary layer (4) curve；The characteristic strips equation of step (4), TG section curves in step (2), obtains the slope in T points downstream；Another T points downstream Slope it is identical with the slope of T points upstream；Obtain the complete curve of contraction section (1)；Step (5) presets and owes expansion modifying factor λ；The outer wall curve of test section (3) is straightway；According to step (2) The characteristic strips equation of middle AD sections of curve obtains the slope of D points upstream；The slope of test section (3) is multiplied by deficient for the slope of D points upstream Expand modifying factor λ；Obtain the complete curve of test section (3).
- 2. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (2) in, the method for building up of coordinate system oxy is：Using T points as coordinate origin, positive direction of the x-axis is direction straight up to be axially directed to outlet extreme direction, y-axis square.
- 3. a kind of supersonic speed integration Nozzle Design method according to claim 2, it is characterised in that：The step (2) in, the characteristic strips equation of AD sections of curves is：The characteristic strips equation of TG sections of curves is：In formula, x is the abscissa of coordinate system oxy；Y is the ordinate of coordinate system oxy；Specific heats of gases ratio in γ jet pipes.
- 4. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (3) in, the method that the momentum thickness δ ' of boundary layer (4) is calculated by toll bar momentum integral equation is：Wherein, H is boundary-layer form factor；CfFor friction coefficient.
- 5. a kind of supersonic speed integration Nozzle Design method according to claim 1, it is characterised in that：The step (3) in, the computational methods of displacement thickness δ are：δ=H δ ' (4).
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Cited By (7)
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| CN109815564A (en) * | 2019-01-09 | 2019-05-28 | 南京航空航天大学 | The supersonic speed propelling nozzle mimetic design method of work off one's feeling vent one's spleen dynamic parameter distribution and determining outlet shapes can be simulated |
| CN110207934A (en) * | 2019-05-28 | 2019-09-06 | 中国航天空气动力技术研究院 | Effectively extend the method for the high enthalpy impulse wind tunnel flowing time of large scale freedom piston |
| CN111220341A (en) * | 2020-01-21 | 2020-06-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Design method of wind tunnel high-Mach-number low-Reynolds-number axisymmetric profile spray pipe |
| CN113946904A (en) * | 2021-08-31 | 2022-01-18 | 中国航天空气动力技术研究院 | Design method of large-size low-noise spray pipe |
| CN114969627A (en) * | 2022-05-08 | 2022-08-30 | 中机新材料研究院(郑州)有限公司 | Method for designing molded surface of supersonic circumferential seam spray pipe for gas atomization powder preparation |
| CN115048730A (en) * | 2022-08-15 | 2022-09-13 | 中国人民解放军国防科技大学 | Axial symmetry supersonic velocity spray pipe optimization design method and device based on grid displacement |
| CN115389155A (en) * | 2022-07-29 | 2022-11-25 | 中国航天空气动力技术研究院 | Design method of hypersonic gas-liquid-solid multiphase flow molded surface spray pipe |
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| CN109815564A (en) * | 2019-01-09 | 2019-05-28 | 南京航空航天大学 | The supersonic speed propelling nozzle mimetic design method of work off one's feeling vent one's spleen dynamic parameter distribution and determining outlet shapes can be simulated |
| CN110207934A (en) * | 2019-05-28 | 2019-09-06 | 中国航天空气动力技术研究院 | Effectively extend the method for the high enthalpy impulse wind tunnel flowing time of large scale freedom piston |
| CN110207934B (en) * | 2019-05-28 | 2021-06-11 | 中国航天空气动力技术研究院 | Method for effectively prolonging flow time of large-size free piston high-enthalpy pulse wind tunnel |
| CN111220341A (en) * | 2020-01-21 | 2020-06-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Design method of wind tunnel high-Mach-number low-Reynolds-number axisymmetric profile spray pipe |
| CN111220341B (en) * | 2020-01-21 | 2020-11-13 | 中国空气动力研究与发展中心超高速空气动力研究所 | Design method of wind tunnel high-Mach-number low-Reynolds-number axisymmetric profile spray pipe |
| CN113946904A (en) * | 2021-08-31 | 2022-01-18 | 中国航天空气动力技术研究院 | Design method of large-size low-noise spray pipe |
| CN113946904B (en) * | 2021-08-31 | 2024-06-11 | 中国航天空气动力技术研究院 | Design method of large-size low-noise spray pipe |
| CN114969627A (en) * | 2022-05-08 | 2022-08-30 | 中机新材料研究院(郑州)有限公司 | Method for designing molded surface of supersonic circumferential seam spray pipe for gas atomization powder preparation |
| CN115389155A (en) * | 2022-07-29 | 2022-11-25 | 中国航天空气动力技术研究院 | Design method of hypersonic gas-liquid-solid multiphase flow molded surface spray pipe |
| CN115389155B (en) * | 2022-07-29 | 2024-08-06 | 中国航天空气动力技术研究院 | Hypersonic gas-liquid-solid multiphase flow molded surface spray pipe design method |
| CN115048730A (en) * | 2022-08-15 | 2022-09-13 | 中国人民解放军国防科技大学 | Axial symmetry supersonic velocity spray pipe optimization design method and device based on grid displacement |
| CN115048730B (en) * | 2022-08-15 | 2022-10-21 | 中国人民解放军国防科技大学 | Axial symmetry supersonic velocity spray pipe optimization design method and device based on grid displacement |
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