CN112035952B - Design method of ejector nozzle experimental device for simulating outflow of aircraft - Google Patents

Design method of ejector nozzle experimental device for simulating outflow of aircraft Download PDF

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CN112035952B
CN112035952B CN202010847794.0A CN202010847794A CN112035952B CN 112035952 B CN112035952 B CN 112035952B CN 202010847794 A CN202010847794 A CN 202010847794A CN 112035952 B CN112035952 B CN 112035952B
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nozzle
outflow
wall surface
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黄河峡
鲁世杰
李子杰
秦源
谭慧俊
林正康
马志明
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Nanjing University of Aeronautics and Astronautics
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    • G06F30/15—Vehicle, aircraft or watercraft design
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    • G06F30/28—Design optimisation, verification or simulation using fluid dynamics, e.g. using Navier-Stokes equations or computational fluid dynamics [CFD]
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Abstract

The invention discloses a design method of an ejector nozzle experimental device for simulating outflow of an aircraft. The method comprises the steps of designing an initial molded line of a meridian plane of an outer flow nozzle under a specific flight Mach number by adopting a characteristic line method, then obtaining a molded line of the meridian plane of the outer flow nozzle which meets requirements by adjusting a symmetry factor G of the outer flow channel and the height h of a throat, further obtaining a three-dimensional model of the outer flow channel, and finally designing an outer flow pressure stabilizing section and a circumferentially distributed air inlet section. The design method utilizes a characteristic line method to design the outer flow channel, can utilize the original experimental assembly to a greater extent, has simple method and easy realization of procedures, and can design the profiles of the outer flow channel under different Mach numbers. The supersonic air flow generated by the design method has no expansion wave and shock wave, has uniform flow parameters, can accurately simulate the outflow environment of the aircraft on the premise of ensuring the economy, and provides a feasible experimental device design method for researching the internal and external flow coupling mechanism of the high-altitude flight working condition of the aircraft.

Description

Design method of ejector nozzle experimental device for simulating outflow of aircraft
Technical Field
The invention relates to the field of aircraft pneumatic experiments, in particular to a design method of an experimental device capable of simulating outflow.
Background
The TBCC nozzle tail pipe works in a large pressure drop ratio range, the change range of the mass flow passing through the TBCC nozzle tail pipe is large, the expansion ratio of the TBCC nozzle tail pipe is changed from 2 under a takeoff state to 15-20 under a supersonic cruise state, a variable-geometry method is adopted to adjust the throat and the expansion angle, and the complexity and the additional resistance of the nozzle pipe configuration are increased. There is therefore a need for a nozzle designed pneumatically to contain multiple flow paths of fluid, which is easily implemented in configuration and does not present additional drag. The jet nozzle used for supersonic aircraft jets the overflow from the boundary layer of the air inlet or the gas (secondary flow) from the cooling flow path and the outer bypass, and mixes the shearing and kinetic energy with the gas (primary flow) flowing out of the main jet pipe, thereby improving the kinetic energy of the secondary flow fluid, and the primary flow fluid and the secondary flow fluid flow out of the jet nozzle together to improve the thrust.
Because the actual ejector nozzle can not be influenced by the outflow of the aircraft in the working process, especially in the low-speed and transonic flight states, the third flow path of the ejector nozzle assists the opening of the inlet valve, the outflow airflow is directly sucked into the ejector nozzle at the moment, and the outflow condition directly influences the working performance of the ejector nozzle.
A common simulation method is to develop a high-speed wind tunnel experiment and simulate high-speed airflow through a wind tunnel Nozzle (Bresnaban D.L, Performance of an aerodyne airborne audience Ejector Nozzle at Mach numbers from 0to 2.0, NASA TM-X-2023). The design method is very complex, a special support system is required to be designed, the primary flow and the secondary flow of the jet nozzle are supplied with air through support, and a special fairing is required to be designed in order to eliminate the interference of the upstream profile of the jet nozzle on a flow field.
Another common lance experiment typically includes a high pressure gas source at the lance inlet to simulate upstream boundary conditions and a low pressure gas source or atmosphere at the lance outlet to simulate downstream conditions. The method can simulate the working state of the spray pipe on the premise of low difficulty and high economy, but cannot accurately simulate the outflow condition, particularly the mutual coupling between the outflow and the main flow and further influences the flow field of the main flow.
Therefore, an actual method of the ejector nozzle experimental device for simulating the outflow of the aircraft needs to be found, the outflow of the aircraft is simulated on the premise of ensuring the economy, and therefore the working condition of the aircraft in the high-altitude actual flight state can be better researched.
Disclosure of Invention
The invention provides a design method of an aircraft outflow simulation experiment device, which can simulate supersonic airflow outside an aircraft on the premise of not carrying out a high-speed wind tunnel experiment.
In order to achieve the purpose, the design method of the invention can adopt the following technical scheme:
a design method of an ejector nozzle experimental device for simulating aircraft outflow comprises the following steps:
(1) providing an experimental device prototype of an aircraft flight Mach number and an injection nozzle, wherein the experimental device prototype comprises a circular rotating flange section (1), a main nozzle section (2) coaxially connected with the rear end of the circular rotating flange section, a secondary flow nozzle section (9) surrounding the main nozzle section and an outer flow nozzle section (8); the outer flow spray pipe section (8) comprises a central body (14) surrounding the secondary flow spray pipe section (9) and an outer cover (15) surrounding the central body; the inner surface of the central body facing the outer cover gradually approaches to one side of the secondary flow spraying pipe section from front to back to form an expansion profile; the inner surface of the outer cover facing the central body gradually expands outwards from front to back, and the length of the inner surface of the outer cover extending backwards exceeds the length of the inner surface of the central body extending backwards;
(2) determining the expansion ratio of the outflow nozzle (8) based on the provided flight Mach number, obtaining an initial molded line (11) of a meridian plane of the outflow nozzle through a characteristic line method, if the molded line meets the geometric constraint and the flow constraint of the existing experimental device, performing the step (6), and if the molded line does not meet the geometric constraint and the flow constraint of the existing experimental device, performing the step (3);
(3) adjusting the asymmetric factor G and the throat height h of the outflow nozzle (8) to enable the meridian surface lower molded line (13) of the outflow nozzle to meet the geometric constraint of the secondary flow nozzle (8) and the flow constraint of an experiment table; the asymmetry factor G of the outflow nozzle (8) is determinedDefined as the initial expansion angle delta of the lower wall surface (13)LInitial expansion angle delta with the upper wall surface (12)UIn a ratio of
Figure GDA0003073071590000021
The value of the asymmetry factor G is between 0 and 1, G-0 indicates that the lower wall surface is along the axial direction, and G-1 indicates that the upper wall surface and the lower wall surface are symmetrical; the lower wall surface (13) is the inner surface of the central body facing the outer cover; the upper wall (12) is the inner surface of the outer cover facing the central body;
(4) when the asymmetry factor G <1, the lower wall surface length is less than the upper wall surface; the expansion wave emitted from the wall surface inlet is reflected on the lower wall surface, the lower wall surface reflection wave is intersected with the expansion wave emitted from the upper wall surface sharp point until the Mach number of a certain reflection wave after the interaction with the last expansion wave of the upper sharp point is equal to the design Mach number, and the flow of the whole core area is determined;
(5) determining upper and lower wall surface points of the outflow nozzle (8); the method comprises the following specific steps that firstly, the upper end point and the lower end point of the throat height h are determined as a point a and a point d, a point b is set, ab is an arc, the center of the ab arc is located on an extension line of da, the radius of the arc is 0.05, and the center angle of the ab arc is the initial expansion angle delta of the upper wall surfaceUThereby determining point b; while being G<1, setting a point e to enable de to be a straight line, determining the point e, continuously calculating reflection and intersection of expansion waves through a characteristic line method until Mach number is designed Mach number after a certain reflected wave is intersected with the last expansion wave of an upper sharp point to obtain a point f, and finally determining upper and lower wall surface points, namely c and f, through wall surface wave elimination to obtain an asymmetric outflow nozzle meridian plane molded line (11);
(6) and (3) rotating the meridian plane line (11) of the outflow nozzle around the axis of the experimental device to obtain a three-dimensional model, and designing an outflow pressure stabilizing section (6) and an outflow distributed air inlet section (7) according to the three-dimensional model.
Furthermore, the height of the outflow pressure stabilizing section (6) is more than 10 times of the height of the throat of the outer runner.
Furthermore, the area of the outflow distributed air inlet section (7) is 1.2 times or more than that of the outflow throat passage
Furthermore, the main flow nozzle (2) and the secondary flow nozzle (5) are determined by the flight Mach number, and a high-quality nozzle internal flow field meeting the experimental requirements is obtained through the experimental device.
Furthermore, the meridian plane line (11) of the outflow nozzle obtained by the characteristic line method is a non-adhesive theoretical design scheme, and boundary layer correction is carried out on the wall surface according to the displacement thickness of the local boundary layer;
the correction formula is as follows:
x＝x0-δ*sinθ
y＝y0+δ*cosθ
in the formula, x0、y0The coordinate of the upper wall surface (12) and the coordinate of the lower wall surface (13) of the outflow nozzle (8) are obtained according to the design of no viscous flow, and theta is the inclination angle of the wall surfaces to the x axis; designing a wall profile of the outflow nozzle (8) according to a non-sticking theory, then calculating the inclination angle of each point of the wall to the x axis, and obtaining the wall coordinate value after the boundary layer is corrected according to the correction formula; delta*For the local boundary layer displacement thickness, the calculation formula is as follows:
δ*＝x·tgβ
β＝a0+a1Ma+a2Ma2+a3Ma3+a4Ma4+a5Ma5
a0＝-7.16665
a1＝6.694431
a2＝-2.209718
a3＝0.3385411
a4＝-0.023611065
a5＝0.00060763751
where β is a characteristic angle whose value is a single-valued function with respect to Mach number Ma, a0、a1、a2、a3、a4And a5Are all empirical values.
Has the advantages that: compared with the prior art, the invention has the following beneficial effects:
the design method skillfully utilizes a characteristic line method to design the outer flow channel, can utilize the original experimental assembly to a greater extent, is simple in design method and easy to realize program, and can obtain the profiles of the outer flow channel under different experimental working conditions only by changing the Mach number of the outer flow. The generated supersonic airflow has no complex expansion waves and shock waves, the flowing parameters are uniform, the outflow of the aircraft can be accurately simulated on the premise of ensuring economy, and a feasible experimental device design method is provided for the research of the high-altitude actual flight condition of the aircraft.
Drawings
FIG. 1 is a schematic diagram of an experimental apparatus for simulating an aircraft outflow ejector nozzle.
Fig. 2 is an enlarged view of a portion of the outflow nozzle of fig. 1.
FIG. 3 is a schematic view of an additional flow nozzle flow field of the pilot nozzle experimental installation simulating aircraft outflow in FIG. 1.
FIG. 4 is a schematic diagram of wall boundary layer correction.
Detailed Description
The invention discloses a design method of an ejector nozzle experimental device capable of simulating outflow of an aircraft. Referring to fig. 1, the prototype of the experimental apparatus includes a circular flange section 1, a main nozzle section 2 coaxially connected to the rear end of the circular flange section, a secondary flow nozzle section 9 surrounding the main nozzle section, and an outer flow nozzle section (8); the outer flow spray pipe section (8) comprises a central body (14) surrounding the secondary flow spray pipe section (9) and an outer cover (15) surrounding the central body; the inner surface of the central body facing the outer cover gradually approaches to one side of the secondary flow spraying pipe section from front to back to form an expansion profile; the inner surface of the outer cover facing the central body gradually expands outwards from front to back, and the length of the inner surface of the outer cover extending backwards exceeds the length of the inner surface of the central body extending backwards. The difficulty of the invention lies in that the existing ejector nozzle experimental device is utilized to design the outflow nozzle section 10 capable of simulating the outflow of an aircraft, and the specific design steps are as follows:
1. providing an experimental device prototype of an aircraft flight Mach number and an injection nozzle, wherein the experimental device prototype comprises a circular-to-circular flange section 1, a main nozzle section 2 coaxially connected with the rear end of the circular-to-circular flange section, a secondary flow nozzle section 9 surrounding the main nozzle section and an outer flow nozzle section 8; the outer flow jet section 8 comprises a central body 14 surrounding the secondary flow jet section 9, an outer cover 15 surrounding the central body; the inner surface of the central body facing the outer cover gradually approaches to one side of the secondary flow spraying pipe section from front to back to form an expansion profile; the inner surface of the outer cover facing the central body gradually expands outwards from front to back, and the length of the inner surface of the outer cover extending backwards exceeds the length of the inner surface of the central body extending backwards.
2. According to flight Mach number simulated by experimental requirements, the expansion ratio of the outer flow nozzle is determined, and the meridian plane molded line of the initial outer flow nozzle is designed by a characteristic line method. For a detailed description and application of the characteristic line method, see "M J levo-Kr, J D Hoffman, Wang Ru Yong, Wu Zong Zhen et al, gas dynamics, Beijing, national defense industry Press, 1984".
3. Considering the assembly problem with the existing model, in order to enable partial parts to be shared as far as possible, the length of the lower wall surface of the external flow passage Laval nozzle is required to be consistent with the length of the injection nozzle model, and the interference of part structures is avoided. And (5) if the molded line meets the geometric constraint and the flow constraint of the existing experimental device, performing step 4 if the molded line does not meet the geometric constraint and the flow constraint of the experimental bench, wherein the lengths of the upper wall surface and the lower wall surface are required to be adjusted to meet the geometric constraint and the flow constraint of the experimental bench.
4. And adjusting the asymmetric factor G and the throat height h of the outflow nozzle to enable the meridian surface lower profile line of the nozzle to meet the integral geometric constraint of the model.
4.1 for an asymmetric nozzle, the upper and lower halves are solved separately due to the asymmetry of the initial expansion region. If the upper wall of the nozzle is longer than the lower wall, the asymmetry factor G defining the nozzle is equal to the initial expansion angle delta of the lower wallLInitial expansion angle delta with upper wallUIn a ratio of
Figure GDA0003073071590000041
The value of the asymmetry factor G is between 0 and 1, G-0 indicates that the lower wall surface is along the axial direction, and G-1 indicates that the upper wall surface and the lower wall surface are symmetrical; the lower wall 13 is the inner surface of the central body facing the outer cover; the upper wall 12 is the inner surface of the cover facing the central body;
4.2 is shown in fig. 2 and 3. The length of the lower wall surface is smaller than that of the upper wall surface, and the expansion wave emitted from the inlet of the upper wall surface is reflected on the lower wall surface. The lower wall surface reflection wave is intersected with the expansion wave emitted after the upper wall surface sharp point until the Mach number of a certain reflection wave after the last expansion wave of the upper sharp point is interacted with the design Mach number (g point in the figure), and all core area flows are determined.
4.3 finally, determining upper and lower wall surface points of the outflow nozzle 8; the method comprises the following steps that firstly, the upper end point and the lower end point of the throat height h are determined as a point a and a point d, a point b is set, ab is an arc, the center of the ab arc is located on an extension line of da, the radius of the arc is 0.05, and the angle of the center of the ab arc is the initial expansion angle delta U of the upper wall surface, so that the point b is determined; meanwhile, when G is less than 1, setting a point e to enable de to be a straight line, determining a point e, continuously calculating reflection and intersection of the expansion waves through a characteristic line method until the Mach number is the designed Mach number after a certain reflected wave is intersected with the last expansion wave of the upper sharp point to obtain a point f, and finally determining upper and lower wall surface points, namely c and f, through wall surface wave elimination to obtain an asymmetric outflow nozzle meridian plane molded line 11;
4.4 please refer to fig. 4, the characteristic line method theory is based on the theoretical design without viscous flow, so that the boundary layer correction needs to be performed on the wall surface, and the correction formula is as follows:
x＝x0-δ*sinθ
y＝y0+δ*cosθ
in the formula, x0、y0Theta is the wall coordinate of the outflow nozzle in a non-viscous flow design, and theta is the angle of inclination of the wall to the x-axis. After the wall molded line of the outflow nozzle is designed according to the inviscid theory, the inclination angle of each point of the wall to the x axis is calculated, and the wall coordinate value after the boundary layer is corrected can be calculated according to the formula. Delta*For local boundary layer displacement thickness, the correction formula is as follows:
x＝x0-δ*sinθ
y＝y0+δ*cosθ
in the formula (I), the compound is shown in the specification,x0、y0the coordinate of the upper wall surface 12 and the coordinate of the lower wall surface 13 of the outflow nozzle 8 are obtained according to the design of no viscous flow, and theta is the inclination angle of the wall surface to the x axis; designing a molded line of the wall surface of the outflow nozzle 8 according to a non-sticking theory, then calculating the inclination angle of each point of the wall surface to the x axis, and obtaining a wall surface coordinate value after the boundary layer is corrected according to the correction formula; delta*For the local boundary layer displacement thickness, the calculation formula is as follows:
δ*＝x·tgβ
β＝a0+a1Ma+a2Ma2+a3Ma3+a4Ma4+a5Ma5
a0＝-7.16665
a1＝6.694431
a2＝-2.209718
a3＝0.3385411
a4＝-0.023611065
a5＝0.00060763751
where β is a characteristic angle whose value is a single-valued function with respect to Mach number Ma, a0、a1、a2、a3、a4And a5Are all empirical values.
5. And rotating the meridian plane profile of the outflow nozzle around the axis of the experimental device to obtain a three-dimensional model, and designing an outflow pressure stabilizing section and a circumferential air inlet structure according to the three-dimensional model. In order to ensure the uniformity of the airflow, the height of the outflow pressure stabilizing section 6 is more than 10 times of the height of the throat of the outer flow channel, and meanwhile, in order to avoid the blockage of the distributed air inlet section 7, the area of the outflow distributed air inlet section 7 needs to be more than 1.2 times of the area of the outflow throat.
In addition, the present invention has many specific implementations and ways, and the above description is only a preferred embodiment of the present invention. It should be noted that, for those skilled in the art, without departing from the principle of the present invention, several improvements and modifications can be made, and these improvements and modifications should also be construed as the protection scope of the present invention.

Claims (4)

1. A design method of an ejector nozzle experimental device for simulating aircraft outflow is characterized by comprising the following steps:
(1) providing an experimental device prototype of an aircraft flight Mach number and an injection nozzle, wherein the experimental device prototype comprises a circular rotating flange section (1), a main nozzle section (2) coaxially connected with the rear end of the circular rotating flange section, a secondary flow nozzle section (9) surrounding the main nozzle section and an outer flow nozzle section (8); the outer flow spray pipe section (8) comprises a central body (14) surrounding the secondary flow spray pipe section (9) and an outer cover (15) surrounding the central body; the secondary flow nozzle section (9) comprises a secondary flow nozzle (5), and a central body (14) surrounds the secondary flow nozzle (5); the inner surface of the central body facing the outer cover gradually approaches to one side of the secondary flow nozzle (5) from front to back to form an expansion profile; the surface of the outer cover facing the central body is gradually expanded outwards from front to back, and the length of the inner surface of the outer cover extending backwards exceeds the length of the inner surface of the central body extending backwards;
(2) determining the expansion ratio of the outflow nozzle (8) based on the provided flight Mach number, obtaining an initial molded line (11) of a meridian plane of the outflow nozzle through a characteristic line method, if the molded line meets the geometric constraint and the flow constraint of the existing experimental device, performing the step (6), and if the molded line does not meet the geometric constraint and the flow constraint of the existing experimental device, performing the step (3);
the meridian plane molded line (11) of the outflow nozzle obtained by the characteristic line method is a non-adhesive theoretical design scheme, and boundary layer correction is carried out on the wall surface according to the displacement thickness of a local boundary layer;
the correction formula is as follows:
x＝x0-δ*sinθ
y＝y0+δ*cosθ
in the formula, x0、y0The coordinate of the upper wall surface (12) and the coordinate of the lower wall surface (13) of the outflow nozzle (8) are obtained according to the design of no viscous flow, and theta is the inclination angle of the wall surfaces to the x axis; designing a wall profile of the outflow nozzle (8) according to a non-sticking theory, then calculating the inclination angle of each point of the wall to the x axis, and obtaining the wall coordinate value after the boundary layer is corrected according to the correction formula; delta*For the local boundary layer displacement thickness, the calculation formula is as follows:
δ*＝x·tgβ
β＝a0+a1Ma+a2Ma2+a3Ma3+a4Ma4+a5Ma5
a0＝-7.16665
a1＝6.694431
a2＝-2.209718
a3＝0.3385411
a4＝-0.023611065
a5＝0.00060763751
where β is a characteristic angle whose value is a single-valued function with respect to Mach number Ma, a0、a1、a2、a3、a4And a5Are all empirical values;
(3) adjusting the asymmetric factor G and the throat height h of the outflow nozzle (8) to ensure that the lower wall surface (13) of the outflow nozzle meets the geometric constraint of the secondary flow nozzle (5) and also meets the flow constraint of an experiment table; the asymmetry factor G of the outflow nozzle (8) is defined as the initial expansion angle delta of the lower wall (13)LInitial expansion angle delta with the upper wall surface (12)UIn a ratio of
Figure FDA0003073071580000021
The value of the asymmetry factor G is between 0 and 1, G-0 indicates that the lower wall surface is along the axial direction, and G-1 indicates that the upper wall surface and the lower wall surface are symmetrical; the lower wall surface (13) is the inner surface of the central body facing the outer cover; the upper wall (12) is the inner surface of the outer cover facing the central body;
(4) when the asymmetry factor G <1, the lower wall surface length is less than the upper wall surface; the expansion wave emitted from the wall surface inlet is reflected on the lower wall surface, the lower wall surface reflection wave is intersected with the expansion wave emitted from the upper wall surface sharp point until the Mach number of a certain reflection wave after the interaction with the last expansion wave of the upper sharp point is equal to the design Mach number, and the flow of the whole core area is determined;
(5) determining upper and lower wall surface points of the outflow nozzle (8); the specific steps are that firstly, the height h of the throat is from top to bottomThe two end points are determined as a point a and a point d, a point b is set, ab is an arc line, the center of the ab arc line is positioned on the extension line of da, the radius of the arc is 0.05, and the center angle of the ab arc line is the initial expansion angle delta of the upper wall surfaceUThereby determining point b; while being G<1, setting a point e to enable de to be a straight line, determining the point e, continuously calculating reflection and intersection of expansion waves through a characteristic line method until Mach number is designed Mach number after a certain reflected wave is intersected with the last expansion wave of an upper sharp point to obtain a point f, and finally determining upper and lower wall surface points, namely c and f, through wall surface wave elimination to obtain an asymmetric outflow nozzle meridian plane molded line (11);
(6) and (3) rotating the meridian plane line (11) of the outflow nozzle around the axis of the experimental device to obtain a three-dimensional model, and designing an outflow pressure stabilizing section (6) and an outflow distributed air inlet section (7) according to the three-dimensional model.
2. The design method of the ejector nozzle experimental device for simulating the outflow of the aircraft according to claim 1, characterized in that: the height of the outflow pressure stabilizing section (6) is more than 10 times of the height of the throat of the outer runner.
3. The design method of the ejector nozzle experimental device for simulating the outflow of the aircraft according to claim 1 or 2, characterized in that: the area of the outflow distributed air inlet section (7) is 1.2 times or more than that of the outflow throat.
4. The design method of the ejector nozzle experimental device for simulating the outflow of the aircraft according to claim 3, characterized in that: the main flow spray pipe (2) and the secondary flow spray pipe (5) are determined by flight Mach number, and a high-quality spray pipe internal flow field meeting experimental requirements is obtained through the experimental device.
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