CN112240831A - Design method of aero-engine intake temperature distortion generator - Google Patents
Design method of aero-engine intake temperature distortion generator Download PDFInfo
- Publication number
- CN112240831A CN112240831A CN202011106676.0A CN202011106676A CN112240831A CN 112240831 A CN112240831 A CN 112240831A CN 202011106676 A CN202011106676 A CN 202011106676A CN 112240831 A CN112240831 A CN 112240831A
- Authority
- CN
- China
- Prior art keywords
- radial
- temperature distortion
- temperature
- distortion generator
- jet
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M15/00—Testing of engines
- G01M15/04—Testing internal-combustion engines
Landscapes
- Chemical & Material Sciences (AREA)
- Engineering & Computer Science (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- General Physics & Mathematics (AREA)
- Testing Of Engines (AREA)
- Aerodynamic Tests, Hydrodynamic Tests, Wind Tunnels, And Water Tanks (AREA)
Abstract
A design method for an aero-engine intake temperature distortion generator belongs to the field of aero-engine test systems and is based on the high-temperature working medium jet principle, a replaceable jet nozzle is provided and used, a high-temperature heat source is injected into an engine channel, the circumferential and radial spatial distribution of a high-temperature area is independently controllable on the premise that the basic structure of the temperature distortion generator is not changed, the temperature distortion generator designed by the method can meet the test requirements of aero-engine temperature distortion in various states, a large amount of test cost is saved, and the test period of aero-engine temperature distortion is shortened.
Description
Technical Field
The invention belongs to the field of aircraft engine test systems, and particularly relates to a design method of an aircraft engine inlet air temperature distortion generator.
Background
In the actual working process of the aircraft engine, the high-temperature fuel gas which is rebounded after the tail flow of a launching weapon, the tail flow of a front aircraft is sucked in the formation flight of a plurality of aircraft, and the tail flow of a vector jet pipe is sucked in the combustion gas or the tail flow of a vertical take-off and landing aircraft is sucked in the vector jet pipe and is jetted to the ground can be generated. After the aircraft engine sucks the high-temperature gas in a short time, the thrust is reduced, the stable working range is reduced, and the engine can be stopped under extreme conditions. The time and space distribution of the temperature of the high-temperature gas sucked into the aircraft engine at the inlet plane is uneven, which is called inlet temperature distortion. In view of the above working conditions, in the development and testing stage of the aircraft engine, the temperature distortion test must be carried out on the engine according to the existing aircraft engine intake temperature distortion evaluation standard, and the main targets of the test include: determining critical parameters of distortion of the engine inlet temperature and evaluating the effectiveness and reliability of the anti-surge system. In the testing stage of the aircraft engine, the air flow with temperature distortion under different working conditions can be artificially produced on a ground test bed and is conveyed to the engine to complete the test. This type of device capable of simulating the fabrication of a flow field with temperature distortion is called a temperature distortion generator. The performance parameters of the temperature distortion generator include: average outlet surface temperature rise, critical temperature rise response, high temperature area range, average surface temperature rise rate, total pressure loss, total pressure distortion and the like. At present, various aero-engine intake temperature distortion generators are developed and put into use in major aviation industries in the world. According to the research of data, the existing temperature distortion generators can be classified into an external heat flow introduction type and an embedded combustion type according to the different heat sources. The first type of external heat flow introduction temperature distortion generator is a flow field which is arranged outside an engine testing channel, generates high-temperature working media in heating, combustion, heat exchange and other modes, then introduces the high-temperature working media into the engine testing channel by using an arranged pipeline, and finally forms a required temperature distortion map or a required temperature distortion coefficient on an engine pneumatic interface. The temperature distortion generator for external heat flow introduction can be divided into the following types according to different working media: high-temperature air generated by the heat exchanger, high-temperature fuel gas generated by external combustion and external high-temperature steam. Typical external heat flow introduction type temperature distortion generators include a multipoint downstream high-temperature jet flow temperature distortion generator developed by the U.S. arnold engineering development center, a high-temperature steam jet flow temperature distortion generator developed by the U.S. naval research institute, and the like. The second type of embedded combustion type temperature distortion generator is characterized in that a micro combustion cavity is designed in an engine air inlet, gaseous or liquid fuel is led in through a pipeline, then generation of different temperature distortion flow fields at an engine inlet is achieved by controlling the combustion state of the micro combustion cavity, or solid fuel is arranged in the combustion cavity of the temperature distortion generator before testing is started, and the fuel is ignited or spontaneously combusted when testing is started to form the temperature distortion flow fields. Typical embedded combustion temperature distortion generators include a hydrogen combustion temperature distortion generator developed by the united states aerospace agency, livis research center, a hydrogen combustion temperature distortion generator developed by the chinese gas turbine institute, and a propane combustion temperature distortion generator developed by the university of virginia.
The two types of temperature distortion generators have advantages and disadvantages respectively, and have different application ranges. The first external heat flow introduction type temperature distortion generator has the advantages that a heat flow generating device is positioned outside an air inlet of an engine, and a heat flow conveying pipeline and a heat flow injection device inside the air inlet are relatively simple in structure. And because the heat flow generating device is positioned outside the simulated pipeline of the engine, the generation of the high-temperature working medium is safer and the operation is easy. The defects of the simulation technology are obvious due to the fact that pipelines are adopted to convey heat flows, the simulation technology is difficult to generate and obtain in a state of large temperature instant change, and the simulation technology is often used in the situation that the surface average temperature rise is relatively low and long-time testing is needed. The second type of embedded combustion type temperature distortion generator can realize the simulation of important temperature distortion parameters such as surface average temperature rise, distortion range and temperature transient, but because the miniature combustion chamber is positioned in the air inlet passage of the engine, the accurate simulation of the temperature distortion parameters needs to independently control the oil supply and flame stability of each combustion chamber, the structure is complex, the control difficulty is high, and the embedded combustion type temperature distortion generator is often used for short-time test with relatively low precision requirements on the circumferential and radial distribution positions of the temperature distortion, but has strict requirements on transient temperature sudden change indexes.
Disclosure of Invention
The invention aims to solve the problems in the prior art, provides a design method of an air inlet temperature distortion generator of an aircraft engine, provides a great improvement on the basis of a first external heat flow introduction type temperature distortion generator, overcomes the defect that the temperature rise spatial distribution of high-temperature jet flow on a pneumatic interface is difficult to accurately control due to the relative fixation of the injection mode and the injection position of a high-temperature jet flow unit, and realizes the accurate and independent controllable temperature flow field distortion generation in the stable circumferential and radial temperature rise ranges.
In order to achieve the purpose, the invention adopts the following technical scheme:
a design method for an aero-engine intake temperature distortion generator comprises the following steps:
(1) determining the radius of a central support ring and the radius of an excircle of an inlet temperature distortion generator according to the maximum radius of an air inlet rectifying cone of an aircraft engine and the radius of an engine inlet, selecting the number of radial supporting rods and the number of jet nozzles on a single radial supporting rod, calculating the minimum fan angle and the minimum annular area, checking whether the minimum fan angle and the minimum annular area meet the design requirements, and if not, reselecting the number of the radial supporting rods and the number of the jet nozzles on the single radial supporting rod;
(2) calculating the radius of the position of the jet nozzle on the single radial strut by using an equal area method;
(3) selecting the section airfoil parameters and the airfoil width of the radial strut;
(4) calculating the position of the maximum divergence angle intersection line according to the position of the jet nozzle at the maximum radius, judging whether the intersection condition is met, returning to the step (1) if the intersection condition is not met, reselecting the number of the radial struts and the number of the jet nozzles on a single radial strut, or returning to the step (3) and reselecting the section airfoil parameters and the width of the radial struts;
(5) calculating and judging the blockage degree, returning to the step (1) if the blockage degree is more than 35%, and reselecting the number of the radial struts and the number of jet nozzles on a single radial strut;
(6) verifying the performance through a scaled model wind tunnel test or a computer numerical simulation, returning to the step (1) if the performance does not meet the design requirement, reselecting the number of the radial struts and the number of jet nozzles on a single radial strut, or returning to the step (3), and reselecting the section airfoil parameters and the airfoil width of the radial struts;
(7) and (5) the performance is verified to be qualified, and the design is completed.
The aero-engine intake temperature distortion generator is arranged in the intake passage and comprises a central support ring, a plurality of radial support rods and an air pipe; the radial strut is connected to the periphery of the central support ring, and the radial strut and the central support ring are completely wrapped inside the air inlet channel; the breather pipe is inserted at the tail end of the radial supporting rod, and a tail end interface of the breather pipe is positioned outside the air inlet channel and connected with the high-temperature working medium generator and used for guiding the high-temperature working medium to the radial supporting rod.
The breather pipe is provided with a jet hole, and the radial strut is provided with a jet nozzle corresponding to the jet hole.
The radial supporting rod is provided with a mounting opening, and the jet nozzle is detachably inserted in the mounting opening.
The jet channel size of the jet nozzle is gradually expanded or arranged in an equal straight way from the inlet to the outlet.
The cross sections of the central support ring and the radial supporting rods are wing-shaped, so that aerodynamic resistance is reduced.
The cross sections of the central support ring and the radial struts have the same airfoil parameters and airfoil widths.
Compared with the prior art, the technical scheme of the invention has the following beneficial effects:
based on the high-temperature working medium jet principle, the invention provides that a replaceable jet nozzle is utilized to inject a high-temperature heat source into an engine channel, the circumferential and radial spatial distribution of a high-temperature area is independently controllable on the premise of not changing the basic structure of the temperature distortion generator, the temperature distortion generator designed by the method can meet the temperature distortion test requirements of the aircraft engine under various states, a large amount of test cost is saved, and the test period of the temperature distortion of the aircraft engine is shortened.
Drawings
FIG. 1 is a schematic view of an aircraft engine inlet temperature distortion test stand;
FIG. 2 is a schematic view of an aircraft engine and air intake;
FIG. 3 is a front view of an aircraft engine inlet in the air intake direction without an air intake temperature distortion generator;
FIG. 4 is a cross-sectional view of the rear intake duct with the intake temperature distortion generator installed;
FIG. 5 is a schematic diagram of an intake air temperature distortion generator;
FIG. 6 is a schematic view of the installation of the radial strut and the jet nozzle of the inlet temperature distortion generator;
FIG. 7 is a reverse flow view of the temperature distortion generator after installation of the jet nozzle;
FIG. 8 is a schematic view of a vent tube;
FIG. 9 is a schematic diagram of a radial strut of an intake temperature distortion generator ejecting high temperature working medium;
FIG. 10 is a schematic view of an inlet temperature distortion generator installed in an inlet duct;
FIG. 11 is a front view in the intake direction of an intake air temperature distortion generator installed in an intake duct;
FIG. 12 is a schematic view of the range of circumferential high temperature zones that can be affected by a single radial strut jet nozzle spray;
FIG. 13 is a schematic view of the range of radial high temperature zones that can be affected by the jet nozzle jet at the same radial location;
FIG. 14 is a schematic view of the gas flow when the single radial strut jet nozzle is fully open;
FIG. 15 is a schematic view of the gas flow when the single radial strut jet nozzle portion is open;
FIG. 16 is a schematic view of a jet nozzle at different divergence angles;
FIG. 17 is a schematic representation of the relative positions of the radial strut cross-section and the aerodynamic interface;
FIG. 18 is a schematic cross-sectional view of a radial strut;
FIG. 19 is a schematic view of the cross-section of adjacent fluidic nozzles at the same radial location circumferentially expanded;
FIG. 20 is a flow chart of a design method of the present invention.
Detailed Description
In order to make the technical problems, technical solutions and advantageous effects of the present invention clearer and clearer, the present invention is further described in detail below with reference to the accompanying drawings and embodiments.
Fig. 1 shows a schematic diagram of an aero-engine intake temperature distortion test stand, (1) is a test stand support, an aero-engine (2) to be tested and an air inlet duct (3) are mounted on a thrust guide rail (4), the thrust guide rail (4) is mounted on the test stand support (1), and in the test process of the aero-engine, the aero-engine (2), the air inlet duct (3) and the thrust guide rail (4) can move back and forth in a small range along the axial direction of the engine, and (5) is the air flow direction sucked into the aero-engine in the test process.
Fig. 2 is a schematic diagram of an aircraft engine (2) and an air inlet (3), the basic structure of the air inlet (3) adopted in the engine ground test is a trumpet-shaped air inlet and an equal straight circular section pipeline, the air inlet (3) is directly butted with the engine (2), and an artificially defined pneumatic interface (6) exists between the air inlet and the engine, as shown by a dotted line in fig. 2. The pneumatic interface (6) defines the interface position of the outlet of the air inlet channel (3) and the inlet of the engine (2), and the pneumatic interface (6) is also the position of the measurement plane of the inlet temperature distortion parameter of the aircraft engine.
FIG. 3 is a front view of an aircraft engine inlet in the air intake direction without an intake temperature distortion generator, (7) is the maximum radius of an engine intake fairing cone, denoted by R1, (8) is the engine inlet radius, denoted by R2, and R1 and R2 are intrinsic parameters of the aircraft engine.
Fig. 4 shows a cross-sectional view of the rear intake duct with the intake temperature distortion generator installed, where the intake temperature distortion generator (9) is installed in the inner equal straight section of the intake duct (3) between the horn intake of the intake duct and the aerodynamic interface (6).
Fig. 5 shows a schematic diagram of an intake air temperature distortion generator, the basic structure of which is composed of a central support ring (10) and a plurality of radial struts (11), the cross sections of the central support ring (10) and the radial struts (11) are airfoil-shaped, so as to reduce aerodynamic drag, and the airfoil parameters and the airfoil widths of the cross sections of the central support ring (10) and the radial struts (11) are the same. A vent pipe (12) is inserted into the tail end of each radial strut (11), and the vent pipe (12) introduces external high-temperature working medium into the radial struts.
Fig. 6 is a schematic view showing the installation of the radial strut and the jet nozzle of the intake temperature distortion generator, wherein the radial strut (11) is a fixed part, and the jet nozzle (13) is a replaceable part which can be replaced according to different test requirements.
Fig. 7 shows a reverse flow view of the temperature distortion generator after the jet nozzle is installed, the jet nozzle (13) is only installed on the radial strut (11), and the jet nozzle is not installed on the central support ring (10).
Fig. 8 is a schematic view of the vent pipe (12), and the position of the jet hole (14) on the vent pipe (12) corresponds to the position of the jet nozzle (13) on the radial strut (11).
Fig. 9 is a schematic diagram showing that the radial strut of the intake temperature distortion generator injects a high-temperature working medium, the high-temperature working medium (15) is generated by an external generator, and may be heated high-temperature air, high-temperature fuel gas or high-temperature steam, the high-temperature working medium (15) is firstly introduced into the radial strut (11) through the vent pipe (12), then enters the jet nozzle (13) through the jet hole (14), and finally is injected into the intake duct to be mixed with air sucked by the engine to generate a temperature distortion flow field.
Fig. 10 shows that the intake temperature distortion generator is installed in the intake duct, after the intake temperature distortion generator (9) is installed inside the intake duct (3), the central support ring (10) and the radial strut (11) are completely wrapped inside the intake duct channel, and the end interface of the vent pipe (12) is located outside the intake duct (3) and is used for connecting the high-temperature working medium generator.
Fig. 11 is a front view of the intake temperature distortion generator installed in the intake duct in the intake direction, (16) is the radius of the center support ring of the intake temperature distortion generator, which is denoted by Rd1, and (17) is the radius of the outer circumference of the intake temperature distortion generator, which is denoted by Rd 2.
FIG. 12 (18) is a plot of the range of circumferential hot zones that can be affected by a single radial strut jet nozzle jet, a parameter that is a design requirement and is expressed in degrees Δ θ.
FIG. 13 is a schematic view showing the range of the radial high-temperature region which can be influenced by the jet nozzle jet at the same radial position, (19) is the radius of the distribution position of the jet nozzles, which is denoted by r (i), and (20) is the radial influence range, which is a parameter of design requirements, which is denoted by Δ S and is expressed in square meters.
Fig. 14 is a schematic view showing the flow of air flow when a single radial strut jet nozzle is fully opened, a high-temperature working medium (15) enters the radial strut (11) through the vent pipe (12), further, the high-temperature working medium (15) enters the jet nozzle (13) from the jet hole (14), is further expanded in the jet nozzle (13) and then is ejected out, mixed with air sucked by the engine, and finally forms a temperature distortion flow field at the pneumatic interface (6), wherein (21) is the central line of the jet nozzle, and (22) is the central line of the engine.
Fig. 15 shows the schematic flow of the gas flow when the jet nozzle part of a single radial strut is opened, under certain test conditions, the jet nozzles (13) on the single radial strut (11) do not need to be fully opened, solid jet nozzles are placed at corresponding positions to block high-temperature working media, and meanwhile, the vent pipes (12) at the positions corresponding to the solid jet nozzles do not have jet holes.
Fig. 16 is a schematic view of the jet nozzles (13) with different divergence angles, the radii of the inlets of the nozzles are the same, the divergence angle of the outlet is between zero and the maximum divergence angle, the maximum divergence angle is denoted by Amax, the divergence angles of the three jet nozzles (13) from left to right in fig. 16 are gradually reduced from Amax to zero, and the solid jet nozzles are special cases of the jet nozzles, play a role in blocking airflow and have no divergence angle parameter.
FIG. 17 is a schematic representation of the relative positions of the radial strut cross-section and the aerodynamic interface, (23) the distance between the airfoil leading edge and the aerodynamic interface (6), indicated by L1, and (23) the fixed design parameters provided by the test requirements.
Fig. 18 shows a schematic cross-sectional view of a radial strut (11), the cross-sectional shape of the radial strut (11) is an airfoil, the basic design parameters are an airfoil and an airfoil width (24), the airfoil width (24) is represented by L2, the airfoil can be obtained by querying an airfoil database of NASA, and 00 series symmetrical airfoils are recommended to reduce aerodynamic drag. (25) The maximum airfoil thickness, designated as H1, is automatically determined after the airfoil parameters are selected. (26) For the radius of the central circle of the radial strut, indicated by R3, and R3 is smaller than the maximum thickness of the airfoil H1(25), R3 is (0.9-0.95) x (H1/2) and R3 is (27) is the distance between the central circle and the tail edge of the airfoil, indicated by L3 and L3 is (0.6-0.65) x L2, in order to ensure the structural strength.
Fig. 19 shows a schematic diagram of the cross-section of adjacent jet nozzles at the same radial position along the circumferential direction, (28) is the inner diameter of the jet nozzle, denoted by Rd3, the inner diameter (28) of the jet nozzle is the same as the radius of the jet hole (14), and is a design parameter, (29) is the outer diameter of the jet nozzle, denoted by Rd4, the diameter of the jet nozzle is the same as the maximum thickness (25) of the airfoil, i.e., Rd4 is H1/2, (30) is the length of the jet nozzle, denoted by Ls, the length (30) of the jet nozzle is the same as the distance between the center circle and the trailing edge of the airfoil, i.e., Ls is L3, (31) is the distance between the jet nozzle and the aerodynamic interface (6), denoted by Ld, and L1-L2, (32) is the distance between the center lines (21) of adjacent jet nozzles at the same radial position along the circumferential direction, denoted by H2, and H2 is 2 × pi r (nr)/Nc, r (Nr) is the maximum radius of the jet nozzle, Nc is the number of radial struts (11), Nr is the number of jet nozzles (13) on a single radial strut, (33) is the maximum divergence angle Amax of the jet nozzle, (34) is the extension of the maximum divergence angle of the nozzle, and (35) is the distance between the intersection of the extension of the divergence angle (34) and the aerodynamic interface (6) at the maximum divergence angle of the jet nozzle and the intersection of the jet nozzle centerline (21) and the aerodynamic interface (6), as indicated by H3.
As shown in fig. 20, the flow of the design method of the distortion generator of the inlet air temperature of the aircraft engine according to the present invention includes the following steps:
the method comprises the following steps: according to the maximum radius R1(7) of an aircraft engine air inlet fairing cone and the radius R2(8) of an engine inlet, determining the radius Rd1(16) of a center support ring of an air inlet temperature distortion generator and the radius Rd2(17) of an outer circle of the air inlet temperature distortion generator, wherein Rd1 is 1.05 × R1, and Rd2 is R2. According to the requirement of testing spatial resolution by intake air temperature distortion, the number Nc of the radial struts (11) and the number Nr of the jet nozzles (13) on a single radial strut are selected, and the minimum circumferential fan angle theta i and the minimum annular area Si are calculated, wherein the theta i is 360/Nc and has the unit of degree, and the Si is pi x (Rd 2)2-Rd12) And (Nr +1), judging whether the minimum circumferential fan-shaped angle theta i is smaller than or equal to a single radial strut circumferential influence range delta theta (18) in the design requirement and judging whether the minimum annular area Si is smaller than or equal to a radial high-temperature area range delta S (20) which can be influenced by the jet nozzle spraying at the same radius position in the design requirement, if both verification judgments are met, performing a step II, and if one verification is not met, reselecting the number Nc of the radial struts (11) and the number Nr of the jet nozzles (13) on the single radial strut.
Step two: calculating the radius r (i) (19) of the distribution position of the jet nozzles on the single radial strut by using an equal area method, wherein
Step three: selecting airfoil parameters of a radial strut section, wherein the airfoil parameters comprise a two-dimensional airfoil abscissa X (X1, X2, X3, … and xN) and an airfoil ordinate Y (Y1, Y2, Y3, … and yN), N is the number of coordinate points, the two-dimensional airfoil parameter coordinates usually adopt NACA4-digit symmetric airfoil series (NACA4-digit symmetric airfoil), the airfoil coordinates can be obtained through published literature query, the queried airfoil abscissa and ordinate are dimensionless parameters, the X coordinate range is 0-1, an airfoil width L2(24) is selected, and the reference airfoil true coordinates X0 and Y0 can be determined, wherein X0 is X L2, Y0 is Y L2, and the reference airfoil true coordinates X0 and Y32 are obtained and then are designed through a computer three-dimensional 3939 0 to obtain the three-dimensional inlet air distortion modeling surface of the generator.
Step four: calculating the position of the intersection line of the maximum divergence angle according to the position of the jet nozzle at the maximum radius, as shown in fig. 19, calculating the distance H3(35) between the intersection point of the extension line (34) of the divergence angle and the pneumatic interface (6) of the jet nozzle under the maximum divergence angle and the intersection point of the central line (21) of the jet nozzle and the pneumatic interface (6) according to four parameters of the inner diameter Rd3(28) of the jet nozzle, the outer diameter Rd4(29) of the jet nozzle, the length Ls (30) of the jet nozzle and the distance Ld (31) between the jet nozzle and the pneumatic interface, and firstly according to the formula:
the maximum divergence angle Amax is calculated and then calculated according to the formula: h3 is Rd3+ tan (Amax) (L1-L2+ L3) to obtain H3(35), the maximum divergence angle extension lines (34) of two adjacent nozzles at the maximum radius need to be intersected to ensure the coverage of the high-temperature jet working medium on all space areas, therefore, the relation between H2 and H3 needs to be judged, when 2 XH 3 is more than or equal to H2, the maximum divergence angle extension lines of the adjacent nozzles can be intersected at the aerodynamic interface, the design is qualified, and if 2 XH 3 is more than or equal to H2<H2, the maximum divergence angle extension lines of adjacent nozzles do not intersect at the aerodynamic interface, the design is unqualified, and the step one is returned to, the number Nc of the radial struts (11) and the number Nr of the jet flow nozzles (13) on a single radial strut are reselected, or the step three is returned to, the airfoil parameters of the cross sections of the radial struts and the airfoil width L2(24) are reselected.
Step five: estimating the clogging degree of the intake temperature distortion generator, wherein the calculation method of the clogging degree comprises the following steps:
if the maximum blockage du is greater than 35%, the method returns to the step one, and the number Nc of the radial struts (11) and the number Nr of the jet nozzles (13) on the single radial strut are reselected.
Step six: if the maximum blockage du is less than or equal to 35%, verifying performance through a scaled model wind tunnel test or a computer numerical simulation, wherein the performance parameters comprise the range of a high-temperature area and the average temperature rise of a surface, and returning to the step I to reselect the quantity Nc of the radial struts (11) and the quantity Nr of jet nozzles (13) on a single radial strut if the performance does not meet the design requirement, or returning to the step III to reselect the section airfoil parameters and the airfoil width L2(24) of the radial struts.
Step seven: and if the verification is qualified, finishing the design.
The working principle of the invention is as follows:
based on the high-temperature working medium jet principle, high-temperature working medium such as high-temperature air, high-temperature fuel gas or high-temperature steam generated outside an engine air inlet passage is guided into an air inlet temperature distortion generator through an air vent pipe, then the high-temperature working medium is sprayed into the air inlet passage by using jet nozzles which are circumferentially and radially distributed on the air inlet temperature distortion generator, and the high-temperature working medium is mixed with normal-temperature air in the air inlet passage, so that the generation of a local high-temperature area is realized, and a temperature distortion flow field required by testing is finally obtained. The invention is characterized in that in different temperature distortion tests, the distribution of high-temperature areas of a temperature distortion flow field is different, the opening and closing of high-temperature jet flow at corresponding spatial positions are realized by utilizing the opening and closing of jet flow holes at different positions on a vent pipe, and the adjustment of the distribution size of high-temperature areas under the condition of jet flow opening is realized by utilizing replaceable jet flow nozzles with different expansion angles. In different test states, the temperature distortion flow fields distributed in different spatial positions can be generated only by replacing the required vent pipes and the jet nozzles.
Claims (7)
1. A design method for an aero-engine intake temperature distortion generator is characterized by comprising the following steps:
(1) determining the radius of a central support ring and the radius of an excircle of an inlet temperature distortion generator according to the maximum radius of an air inlet rectifying cone of an aircraft engine and the radius of an engine inlet, selecting the number of radial supporting rods and the number of jet nozzles on a single radial supporting rod, calculating the minimum fan angle and the minimum annular area, checking whether the minimum fan angle and the minimum annular area meet the design requirements, and if not, reselecting the number of the radial supporting rods and the number of the jet nozzles on the single radial supporting rod;
(2) calculating the radius of the position of the jet nozzle on the single radial strut by using an equal area method;
(3) selecting the section airfoil parameters and the airfoil width of the radial strut;
(4) calculating the position of the maximum divergence angle intersection line according to the position of the jet nozzle at the maximum radius, judging whether the intersection condition is met, returning to the step (1) if the intersection condition is not met, reselecting the number of the radial struts and the number of the jet nozzles on a single radial strut, or returning to the step (3) and reselecting the section airfoil parameters and the width of the radial struts;
(5) calculating and judging the blockage degree, returning to the step (1) if the blockage degree is more than 35%, and reselecting the number of the radial struts and the number of jet nozzles on a single radial strut;
(6) verifying the performance through a scaled model wind tunnel test or a computer numerical simulation, returning to the step (1) if the performance does not meet the design requirement, reselecting the number of the radial struts and the number of jet nozzles on a single radial strut, or returning to the step (3), and reselecting the section airfoil parameters and the airfoil width of the radial struts;
(7) and (5) the performance is verified to be qualified, and the design is completed.
2. The utility model provides an aeroengine inlet air temperature distortion generator which characterized in that: the aero-engine intake temperature distortion generator is arranged in the intake passage and comprises a central support ring, a plurality of radial support rods and an air pipe; the radial strut is connected to the periphery of the central support ring, and the radial strut and the central support ring are completely wrapped inside the air inlet channel; the breather pipe is inserted at the tail end of the radial supporting rod, and a tail end interface of the breather pipe is positioned outside the air inlet channel and connected with the high-temperature working medium generator and used for guiding the high-temperature working medium to the radial supporting rod.
3. An aircraft engine inlet air temperature distortion generator as claimed in claim 2, wherein: the breather pipe is provided with a jet hole, and the radial strut is provided with a jet nozzle corresponding to the jet hole.
4. An aircraft engine inlet air temperature distortion generator as claimed in claim 3, wherein: the radial supporting rod is provided with a mounting opening, and the jet nozzle is detachably inserted in the mounting opening.
5. An aircraft engine inlet air temperature distortion generator as claimed in claim 3, wherein: the jet channel size of the jet nozzle is gradually expanded or arranged in an equal straight way from the inlet to the outlet.
6. An aircraft engine inlet air temperature distortion generator as claimed in claim 2, wherein: the cross sections of the central support ring and the radial supporting rods are wing-shaped, so that aerodynamic resistance is reduced.
7. An aircraft engine inlet air temperature distortion generator as claimed in claim 6, wherein: the cross sections of the central support ring and the radial struts have the same airfoil parameters and airfoil widths.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011106676.0A CN112240831B (en) | 2020-10-15 | 2020-10-15 | Design method of aero-engine intake temperature distortion generator |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011106676.0A CN112240831B (en) | 2020-10-15 | 2020-10-15 | Design method of aero-engine intake temperature distortion generator |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112240831A true CN112240831A (en) | 2021-01-19 |
| CN112240831B CN112240831B (en) | 2021-06-29 |
Family
ID=74169239
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011106676.0A Active CN112240831B (en) | 2020-10-15 | 2020-10-15 | Design method of aero-engine intake temperature distortion generator |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112240831B (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113418713A (en) * | 2021-06-21 | 2021-09-21 | 中国航发沈阳发动机研究所 | Combined distortion generator of engine |
| CN114658566A (en) * | 2022-03-06 | 2022-06-24 | 西北工业大学 | Temperature distortion generator based on pulse detonation combustion and test method thereof |
| CN114876637A (en) * | 2022-04-13 | 2022-08-09 | 太仓点石航空动力有限公司 | Unsteady load engine inlet total pressure distortion simulation device, method and system |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104848904A (en) * | 2015-06-05 | 2015-08-19 | 中国航天空气动力技术研究院 | Air duct flow measuring system |
| CN106153346A (en) * | 2016-07-20 | 2016-11-23 | 南京航天航空大学 | A kind of inflatable vane type swirl flow distortion generator and method for generation |
| CN106457655A (en) * | 2014-05-29 | 2017-02-22 | 帝斯克玛股份有限公司 | Injection nozzle with finish cooling and counter pressure |
| CN108132149A (en) * | 2017-11-29 | 2018-06-08 | 中国航发沈阳黎明航空发动机有限责任公司 | A kind of aero-engine inlet distortion wire examination method |
| CN108445309A (en) * | 2018-06-05 | 2018-08-24 | 沈阳工业大学 | A kind of temperature gradient disc insulator surface charge measuring device and method off field |
| US20190323909A1 (en) * | 2018-04-24 | 2019-10-24 | The Boeing Campany | Powered total pressure measurement rake with telemetry |
| CN111175053A (en) * | 2020-01-03 | 2020-05-19 | 北京航空航天大学 | Air inlet distortion simulation device for aircraft engine |
-
2020
- 2020-10-15 CN CN202011106676.0A patent/CN112240831B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106457655A (en) * | 2014-05-29 | 2017-02-22 | 帝斯克玛股份有限公司 | Injection nozzle with finish cooling and counter pressure |
| CN104848904A (en) * | 2015-06-05 | 2015-08-19 | 中国航天空气动力技术研究院 | Air duct flow measuring system |
| CN106153346A (en) * | 2016-07-20 | 2016-11-23 | 南京航天航空大学 | A kind of inflatable vane type swirl flow distortion generator and method for generation |
| CN108132149A (en) * | 2017-11-29 | 2018-06-08 | 中国航发沈阳黎明航空发动机有限责任公司 | A kind of aero-engine inlet distortion wire examination method |
| US20190323909A1 (en) * | 2018-04-24 | 2019-10-24 | The Boeing Campany | Powered total pressure measurement rake with telemetry |
| CN108445309A (en) * | 2018-06-05 | 2018-08-24 | 沈阳工业大学 | A kind of temperature gradient disc insulator surface charge measuring device and method off field |
| CN111175053A (en) * | 2020-01-03 | 2020-05-19 | 北京航空航天大学 | Air inlet distortion simulation device for aircraft engine |
Non-Patent Citations (2)
| Title |
|---|
| M. W. MCLLVEEN: "Further test results with the airjet distortion generator system", 《AIAA/SAE/ASME 15TH JOINT PROPULSION CONFERENCE》 * |
| 姜健等: "进气道/发动机相容性评价体系的完善与发展", 《科学技术与工程》 * |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113418713A (en) * | 2021-06-21 | 2021-09-21 | 中国航发沈阳发动机研究所 | Combined distortion generator of engine |
| CN114658566A (en) * | 2022-03-06 | 2022-06-24 | 西北工业大学 | Temperature distortion generator based on pulse detonation combustion and test method thereof |
| CN114658566B (en) * | 2022-03-06 | 2024-01-30 | 西北工业大学 | Temperature distortion generator based on pulse detonation combustion and test method thereof |
| CN114876637A (en) * | 2022-04-13 | 2022-08-09 | 太仓点石航空动力有限公司 | Unsteady load engine inlet total pressure distortion simulation device, method and system |
Also Published As
| Publication number | Publication date |
|---|---|
| CN112240831B (en) | 2021-06-29 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN112240831B (en) | Design method of aero-engine intake temperature distortion generator | |
| Schmid et al. | The impact of realistic inlet swirl in a 1 ½ stage axial turbine | |
| CN111649948A (en) | Air inlet system for annular blade cascade performance tester | |
| Krichbaum et al. | A large scale turbine test rig for the investigation of high pressure turbine aerodynamics and heat transfer with variable inflow conditions | |
| CN112229639B (en) | Design method of aero-engine intake total pressure distortion generation device | |
| CN115127822A (en) | Air inlet channel experimental device | |
| CN107891970A (en) | The active thermal protection system of hypersonic aircraft gaseous film control | |
| Fu et al. | Characteristics of the swirling flow generated by an axial swirler | |
| Hall et al. | Design of a non-reacting combustor simulator with swirl and temperature distortion with experimental validation | |
| Ogorodnikov et al. | Russian research on experimental hydrogen-fueled dual-mode scramjet: conception and preflight tests | |
| Wang et al. | Influence of sleeve angle on the LBO performance of TeLESS-II combustor | |
| Hall et al. | Experimental Study of Non-Reacting Low NOx Combustor Simulator for Scaled Turbine Experiments | |
| Kao et al. | Effect of dome geometry on swirling flow field characteristics of a counter-rotating radial-radial swirler | |
| Carrotte et al. | Detailed measurements on a modern combustor dump diffuser system | |
| Kilik | Better swirl generation by using curved vane swirlers | |
| Kim et al. | Effect of splitter location on the characteristics of a vane-type pre-swirl system | |
| CN112179605B (en) | Ejector nozzle experimental device for simulating outflow of aircraft | |
| Shakariyants et al. | Generic geometry definition of the aircraft engine combustion chamber | |
| Nichols et al. | Aerodynamic control of an inlet flow in crosswind using peripheral bleed actuation | |
| Walker et al. | Impact of a cooled cooling air installation on the external aerodynamics of a gas turbine combustion system | |
| Rasimarzabadi et al. | Evaluation of Inlet Conditions of an AFT Mounted BLI Design in a Sub-Scale Test Rig | |
| Walker et al. | Experimental study of the unsteady aerodynamics the compressor-combustor interface of a lean burn combustion system | |
| Walker et al. | Compressor/diffuser/combustor aerodynamic interactions in lean module combustors | |
| CN114658566B (en) | Temperature distortion generator based on pulse detonation combustion and test method thereof | |
| CN213148312U (en) | Annular cascade performance tester |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |