Aero-engine whole machine fulcrum stiffness test device and method and test design method

Abstract

The application provides an aero-engine whole machine fulcrum stiffness test device, method and test design method, and the test device comprises a whole machine fulcrum loading test tool, and the whole machine fulcrum loading test tool comprises a constraint tool, a fulcrum displacement measuring device and a fulcrum loading tool; the constraint tool is used for fixing a test aero-engine and comprises a fixing assembly and a hoisting device; the fixing assembly is arranged around the test aero-engine, and the hoisting device is connected and fixed to the test aero-engine; the fulcrum displacement measuring device comprises a support frame and a supporting connection structure, the supporting connection structure is arranged through the inside of the test aero-engine, and the supporting connection structure provides a plurality of force applying fulcrums for the test aero-engine; the support frame is used for supporting the supporting connection structure; the fulcrum loading tool is arranged on one side of the test aero-engine in the axial direction, the fulcrum loading tool extends into and is connected to the inside of the supporting connection structure along the axial direction of the supporting connection structure, and the fulcrum loading tool is used for applying a radial force to the supporting connection structure.

Description

Technical Field

[0001] This invention relates to the field of aero-engine stiffness testing technology, specifically to an aero-engine overall support stiffness testing device, method, and test design method. Background Technology

[0002] The overall support stiffness of an aero-engine has a significant impact on its safety and reliability. The magnitude of the support stiffness directly affects the analysis and calculation of the rotor's critical speed and is a crucial parameter in engine rotor dynamics design. Currently published literature includes experimental studies on the overall engine casing flexibility and squirrel cage spring support stiffness to verify the overall engine support stiffness design; however, no literature has been found that comprehensively verifies the stiffness of all support points. The challenges of testing the overall engine support stiffness are as follows:

[0003] I. The overall support stiffness test of an aero-engine involves various aspects such as the installation constraint design, loading design, measurement design, and data processing of the entire engine. Therefore, determining the correct test principle and designing a complete set of overall support stiffness test methods is quite complex.

[0004] Second, aero engines have large axial dimensions, and some pivot points are designed inside the engine, far from the air intake or exhaust nozzle. These pivot points are located in confined spaces, surrounded by the engine casing and support plates, making it impossible to directly measure them using traditional methods or indirectly measure their displacement using rigid rods (cylinders). Testing pivot point displacement requires a separate test bracket designed and installed inside the engine casing. The displacement measurement points and the test bracket must be positioned synchronously, making the bracket structure difficult to design and install.

[0005] Third, for the support points of the cartridge support structure on the whole machine, not only the rigidity of the receiver is required, but also the rigidity of the cartridge support.

[0006] In view of this, the inventors of this application have designed a test device, method and test design method for the overall support stiffness of an aero-engine, in order to overcome the above-mentioned technical problems. Summary of the Invention

[0007] The technical problem to be solved by the present invention is to overcome the shortcomings of the existing technology in the test of fully verifying the stiffness of all support points of the whole engine, and to provide a test device, method and test design method for the stiffness of the support points of the whole engine.

[0008] The present invention solves the above-mentioned technical problems through the following technical solution:

[0009] This invention provides a test apparatus for the overall support stiffness of an aero-engine, characterized in that the test apparatus includes an overall support loading test fixture, which comprises: a constraint fixture, a support displacement measuring device, and a support loading fixture; the constraint fixture is used to fix the aero-engine under test, and includes a fixing component and a hoisting device; the fixing component surrounds the aero-engine under test, and the hoisting device is connected and fixed to the aero-engine under test; the support displacement measuring device includes a support frame and a support connection structure, the support connection structure being disposed through the interior of the aero-engine under test, providing multiple force application points for the aero-engine under test; the support frame is used to support the support connection structure; the support loading fixture is disposed on one side of the axial direction of the aero-engine under test, the support loading fixture extends along the axial direction of the support connection structure and is connected to the interior of the support connection structure, and the support loading fixture is used to apply a radial force to the support connection structure.

[0010] According to one or more embodiments of the present invention, the aero-engine whole-machine support stiffness testing device further includes a projectile stiffness testing fixture, which can simulate the actual force on the projectile and can continuously apply force to the projectile; the projectile stiffness testing fixture includes a projectile structure, a projectile loading plate is provided on the side of the projectile structure, a first loading rod is provided through the center of the projectile structure and the projectile loading plate, the two ends of the first loading rod are connected to one end of a connector, the other end of the connector is connected to the two ends of a second loading rod, and a loading actuator is provided in the middle of the second loading rod.

[0011] According to one or more embodiments of the present invention, the fixing assembly includes at least two columns and a crossbeam, the hoisting device includes a hanging plate, the two columns are fixedly disposed on both radial sides of the test aircraft engine, the two ends of the crossbeam are connected to the upper ends of the two columns, the crossbeam is disposed on the upper part of the test aircraft engine, the hanging plate is disposed on the lower part of the crossbeam, and the hanging plate is connected to the upper part of the test aircraft engine.

[0012] According to one or more embodiments of the present invention, the column includes at least two front columns and at least two rear columns, the crossbeam includes a front crossbeam and a rear crossbeam, and the mounting plate includes a front mounting section mounting plate and a rear mounting section mounting plate; both ends of the front crossbeam are connected to the upper ends of the two front columns, both ends of the rear crossbeam are connected to the upper ends of the two rear columns, a longitudinal beam connects the front columns and the rear columns, a front mounting section mounting plate is provided at the lower part of the front crossbeam, the front mounting section mounting plate is connected to the upper part of the front mounting section of the tested aero-engine, and a rear mounting section mounting plate is provided at the lower part of the rear crossbeam, the rear mounting section mounting plate is connected to the upper part of the rear mounting section of the tested aero-engine.

[0013] According to one or more embodiments of the present invention, the support connection structure includes a front support, an inner cylinder, and a rear support connected in sequence; the support frame includes a front support frame and a rear support frame; the front support frame is connected to the axial end of the front support, and the rear support frame is connected to the axial end of the rear support.

[0014] According to one or more embodiments of the present invention, the front support includes an annular first support plate, the inner ring of the first support plate being connected to one side of a connecting cylinder, the other side of the connecting cylinder being connected to the outer ring of an annular second support plate, the second support plate being connected to a plurality of spoke-shaped third support plates, and the plurality of third support plates being connected to the inner cylinder; a support cylinder is provided at the upper end of the front support frame, and the front support frame is connected to the inner ring of the second support plate through the support cylinder.

[0015] According to one or more embodiments of the present invention, the rear support includes a plurality of fourth support plates, one end of the fourth support plate being vertically connected to the inner cylinder, the other end of the fourth support plate being vertically connected to one end of the first connecting plate, the other end of the first connecting plate being vertically connected to one end of the fifth support plate, the other end of the fifth support plate being vertically connected to one end of the second connecting plate, the other end of the second connecting plate being vertically connected to the outer ring of the annular sixth support plate, and a plurality of third connecting plates being provided on the outer edge of the outer ring of the sixth support plate; a support cylinder is provided at the upper end of the rear support frame, and the rear support frame is connected to the inner ring of the sixth support plate through the support cylinder.

[0016] According to one or more embodiments of the present invention, the fulcrum loading fixture is disposed on one side of the test aircraft engine along the axial direction, the fulcrum loading fixture includes an actuating component, a loading lever, and a support point assembly; a first end of the loading lever passes through and is connected to the interior of the test aircraft engine along the axial direction; the actuating component is connected to a second end of the loading lever, and the actuating component is used to apply force to the second end of the loading lever; the support point assembly is connected to the middle of the loading lever.

[0017] According to one or more embodiments of the present invention, the actuation assembly includes an actuation device, a loading end force sensor, and a loading end bearing assembly; the actuation device is connected to the loading end force sensor, the loading end force sensor is connected to the loading end bearing assembly, and the loading end bearing assembly is connected to the second end of the loading lever.

[0018] According to one or more embodiments of the present invention, the support point assembly includes a support point pad, a support point bearing assembly, and a support point force sensor; the support point pad is connected to the support point bearing assembly, the support point bearing assembly is connected to the support point force sensor, and the support point force sensor is connected to the middle of the loading lever.

[0019] According to one or more embodiments of the present invention, a thrust tie rod assembly is provided between the front mounting section and the rear mounting section of the tested aero-engine.

[0020] This invention also provides a design method for testing the overall support stiffness of an aero-engine. The method is characterized in that it is used to design the aero-engine overall support stiffness testing device as described above. The design method includes the following steps: S1, designing constraint fixtures based on the actual hoisting state of the aero-engine under test; S2, designing support loading fixtures based on the structural characteristics and loading requirements of the overall support of the aero-engine under test; S3, designing a support displacement measuring device based on the support displacement measurement requirements and the test installation state; S4, designing spring support stiffness testing fixtures; S5, constructing a system of aero-engine overall support stiffness testing methods.

[0021] This invention also provides a method for testing the overall support stiffness of an aero-engine, characterized in that the method uses the aero-engine overall support stiffness testing device described above, and the method includes the following steps: P1, the aero-engine under test is installed and connected to the overall support loading test fixture; P2, a support displacement measurement test is performed on the overall aero-engine under test, measuring the displacement at each support point, recording the relationship between the displacement and the load, and obtaining the stiffness of each support point by linear fitting the data; P3, the actual stress on the support structure of the aero-engine under test is simulated by the support stiffness test fixture, and the support structure is continuously loaded to perform the support stiffness test, recording the relationship between the load and the displacement, and obtaining the support stiffness by linear fitting the data.

[0022] The positive and progressive effects of this invention are as follows:

[0023] The present invention provides an aero-engine overall pivot stiffness testing device, method, and test design method that have at least the following advantages:

[0024] This invention relates to an aero-engine overall support stiffness testing device, method, and test design method, which integrates overall engine installation constraint design, loading design, measurement design, and data processing. It can comprehensively and systematically realize the stiffness test of each support point of the entire engine, and provide test support for the rationality of the finite element modeling of the entire engine. Attached Figure Description

[0025] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, in which the same reference numerals always denote the same features, wherein:

[0026] Figure 1 This is a simplified structural diagram of an aero-engine and a schematic diagram of the distribution of pivot bearings.

[0027] Figure 2a This is a three-dimensional schematic diagram of the installation of the aero-engine whole-machine support stiffness testing device of the present invention.

[0028] Figure 2b This is a schematic cross-sectional view of the installation of the aero-engine whole-machine support stiffness testing device of the present invention.

[0029] Figure 3 This is a schematic diagram of the constraint fixture in the aero-engine whole-machine support stiffness testing device of the present invention.

[0030] Figure 4 This is a schematic diagram of the support loading fixture in the aero-engine whole-engine support stiffness testing device of the present invention.

[0031] Figure 5 This is a schematic diagram illustrating the measurement principle of the second fulcrum in the aero-engine whole-machine fulcrum stiffness testing device and method of the present invention.

[0032] Figure 6 This is a schematic diagram of the support displacement measuring device in the aero-engine whole-machine support stiffness testing device of the present invention.

[0033] Figure 7 This is a schematic diagram of the structure of the spring support stiffness testing fixture in the aero-engine whole-machine support stiffness testing device of the present invention.

[0034] Figure 8 This is a flowchart illustrating the test method for the overall support stiffness of an aero-engine according to the present invention.

[0035] [Attached image labels]

[0036] 1000 whole machine fulcrum loading test fixture

[0037] Constraint fixture 1100

[0038] Fixed component 1110

[0039] Column 1111

[0040] Front pillar 1111a

[0041] Rear pillar 1111b

[0042] 1112 crossbeam

[0043] Front crossbeam 1112a

[0044] Rear crossbeam 1112b

[0045] Longitudinal beam 1113

[0046] Inner triangle support 1114

[0047] Outer triangle support 1115

[0048] Lifting device 1120

[0049] Front mounting plate 1121

[0050] Rear installation section hanging plate 1122

[0051] 1200 fulcrum displacement measuring device

[0052] Support frame 1210

[0053] Front support frame 1211

[0054] Rear support frame 1212

[0055] Support cylinder 1213

[0056] Support connection structure 1220

[0057] Front support 1230

[0058] First board 1231

[0059] Connecting cylinder 1232

[0060] Second board 1233

[0061] The third board 1234

[0062] Inner cylinder 1240

[0063] Rear support 1250

[0064] Fourth board 1251

[0065] First connecting plate 1252

[0066] Fifth board 1253

[0067] Second connecting plate 1254

[0068] Sixth board 1255

[0069] Third connecting plate 1256

[0070] Pivot loading fixture 1300

[0071] Actuator 1310

[0072] Actuator 1311

[0073] Loading end force sensor 1312

[0074] Loading end bearing assembly 1313

[0075] Loading lever 1320

[0076] Load the first end of the lever 1321

[0077] Load the second end of the lever 1322

[0078] Pivot loading disk 1323

[0079] Support point component 1330

[0080] Support point pad 1331

[0081] Support point bearing assembly 1332

[0082] Support point force sensor 1333

[0083] 2000 Spear Stiffness Testing Fixture

[0084] 2100 missile support structure

[0085] 2200 ammunition loading disk

[0086] First loading rod 2300

[0087] Connector 2400

[0088] Second loading rod 2500

[0089] Loading actuator 2600

[0090] Test aircraft engine 3000

[0091] First fulcrum 3100

[0092] Second fulcrum 3200

[0093] Third fulcrum 3300

[0094] Fourth fulcrum 3400

[0095] Fifth fulcrum 3500

[0096] Front mounting section suspension assembly 3600

[0097] Rear-mounted suspension assembly 3700

[0098] Thrust-pull rod assembly 3800 Detailed Implementation

[0099] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0100] Embodiments of the invention will now be described in detail with reference to the accompanying drawings. Preferred embodiments of the invention will now be described in detail, examples of which are illustrated in the drawings. Wherever possible, the same reference numerals will be used in all the drawings to denote the same or similar parts. Furthermore, although the terminology used herein is selected from commonly known and used terminology, some terms mentioned in this specification may have been chosen by the applicant at his or her discretion, and their detailed meanings are explained in the relevant sections of the description herein. Moreover, the invention should be understood not only by the actual terms used, but also by the meaning implied by each term.

[0101] like Figures 1-6 As shown, the present invention provides a test device for the overall support stiffness of an aero-engine. The test device for the overall support stiffness of an aero-engine includes an overall support loading test fixture 1000, which includes: a constraint fixture 1100, a support displacement measuring device 1200, and a support loading fixture 1300.

[0102] The constraint fixture 1100 is used to fix the test aircraft engine 3000. The constraint fixture 1100 includes a fixing component 1110 and a hoisting device 1120. The fixing component 1110 surrounds the test aircraft engine 3000, and the hoisting device 1120 is connected and fixed to the test aircraft engine 3000.

[0103] The fulcrum displacement measuring device 1200 includes a support frame 1210 and a support connection structure 1220. The support connection structure 1220 is disposed inside the test aero-engine 3000 and provides multiple force application fulcrums for the test aero-engine 3000. The support frame 1210 is used to support the support connection structure 1220.

[0104] The fulcrum loading fixture 1300 is disposed on one side of the axial direction of the tested aero-engine 3000. The fulcrum loading fixture 1300 extends into and connects to the interior of the support connection structure 1220 along the axial direction of the support connection structure 1220. The fulcrum loading fixture 1300 is used to apply a radial force to the support connection structure 1220.

[0105] Taking the overall support stiffness test of a certain type of aero-engine as an example, the test objects are all stator casings, bearing housings, and mounting sections of the tested aero-engine 3000. The structural form and the distribution of bearings at each support point are as follows: Figure 1 As shown. Figure 1Pivot bearings are installed at the first fulcrum 3100, the second fulcrum 3200, the third fulcrum 3300, the fourth fulcrum 3400, and the fifth fulcrum 3500. The test aircraft engine 3000 is also equipped with a front mounting section sling assembly 3600 and a rear mounting section sling assembly 3700 to connect with the hoisting device 1120, thereby fixing the test aircraft engine 3000 to the constraint fixture 1100.

[0106] Figure 2a and Figure 2b This is the overall assembly drawing of the aero-engine whole-engine support stiffness testing device of the present invention for conducting aero-engine whole-engine support stiffness tests.

[0107] The principle of the test using the aero-engine whole-machine support stiffness test device of the present invention is to simulate the installation state of the engine during actual operation. Radial loads are applied to each support point by lever loading. The relationship between the displacement of each support point and the load is recorded by measuring the displacement at the support point. The stiffness of each support point is obtained after linear fitting.

[0108] The present invention provides a test device for the stiffness of the support points of an aero-engine, which solves the problem that the displacement of the support points inside the aero-engine is difficult to measure directly. By installing the support connection structure 1220 of the support displacement measuring device 1200 inside the aero-engine, the displacement measuring point and the support connection structure 1220 are positioned synchronously, thus solving the problems of difficult test support structure design and installation.

[0109] like Figure 7 As shown, in a preferred embodiment of the aero-engine whole-engine support stiffness test device of the present invention, the aero-engine whole-engine support stiffness test device further includes a projectile stiffness test fixture 2000, which can simulate the actual force on the projectile and can continuously apply force to the projectile.

[0110] The spring support stiffness testing fixture 2000 includes a spring support structure 2100. A spring support loading plate 2200 is provided on the side of the spring support structure 2100. A first loading rod 2300 is passed through the center of the spring support structure 2100 and the spring support loading plate 2200. The two ends of the first loading rod 2300 are connected to one end of a connector 2400. The other end of the connector 2400 is connected to both ends of a second loading rod 2500. A loading actuator 2600 is provided in the middle of the second loading rod 2500.

[0111] It should be noted that, in order to measure the stiffness of the spring support, this invention designs a spring support stiffness testing fixture 2000 that can simulate the actual force on the spring support and achieve continuous loading. Preferably, the spring support structure 2100 is constrained to a rigid mounting plate. The spring support loading disk 2200 simulates the contact between the bearing and the spring support structure 2100 and is axially positioned within the spring support. A first loading rod 2300 is inserted through the center of the spring support loading disk 2200. Connecting parts 2400 are symmetrically distributed at both ends of the first loading rod 2300 with the axial midpoint of the spring support loading disk 2200 as the center. The connecting parts 2400 are connected to the upper second loading rod 2500. A loading actuator 2600 is set at the middle position of the upper second loading rod 2500 to achieve continuous loading. Preferably, displacement sensors are arranged as needed on the outer surface of the spring support structure 2100 to record the relationship between the load and each displacement. The spring support stiffness can be obtained by linearly fitting the data.

[0112] The missile support stiffness test fixture 2000 simulates the structure of the missile support in series at the support point of the entire aero-engine. In order to accurately obtain the stiffness at the rotor support, it is necessary to further obtain the stiffness of the missile support structure 2100 and calculate the complete stiffness at the support point using the following formula (1).

[0113]

[0114] Wherein, K is the complete stiffness at the fulcrum, K1 is the stiffness from the installation point of the elastic support structure 2100 to the constraint point of the installation section, and K2 is the stiffness of the cascaded elastic support at the fulcrum.

[0115] The present invention provides an aero-engine overall support stiffness testing device that takes into account the support points of the tandem support structure 2100 on the aero-engine overall and uses a support stiffness testing fixture 2000 to measure the support stiffness.

[0116] like Figure 3 As shown, in a preferred embodiment of the aero-engine overall support stiffness testing device of the present invention, the fixing component 1110 includes at least two columns 1111 and a crossbeam 1112, and the hoisting device 1120 includes a hanging plate. The two columns 1111 are fixedly arranged on both radial sides of the aero-engine under test 3000. The two ends of the crossbeam 1112 are connected to the upper ends of the two columns 1111. The crossbeam 1112 is arranged on the upper part of the aero-engine under test 3000, and the hanging plate is arranged on the lower part of the crossbeam 1112. The hanging plate is connected to the upper part of the aero-engine under test 3000.

[0117] like Figure 3As shown, in a preferred embodiment of the aero-engine whole-machine support stiffness test device of the present invention, the column 1111 includes at least two front columns 1111a and at least two rear columns 1111b, the crossbeam 1112 includes a front crossbeam 1112a and a rear crossbeam 1112b, and the hanging plate includes a front mounting section hanging plate 1121 and a rear mounting section hanging plate 1122;

[0118] The two ends of the front crossbeam 1112a are connected to the upper ends of the two front columns 1111a, and the two ends of the rear crossbeam 1112b are connected to the upper ends of the two rear columns 1111b. A longitudinal beam 1113 connects the front columns 1111a and the rear columns 1111b. A front mounting section hanger plate 1121 is provided at the lower part of the front crossbeam 1112a. The front mounting section hanger plate 1121 is connected to the upper part of the front mounting section of the tested aircraft engine 3000. A rear mounting section hanger plate 1122 is provided at the lower part of the rear crossbeam 1112b. The rear mounting section hanger plate 1122 is connected to the upper part of the rear mounting section of the tested aircraft engine 3000.

[0119] The constraint fixture 1100 is used for the installation constraint of the tested aero-engine 3000, and the installation constraint method adopts a horizontal hoisting form. Preferably, the constraint fixture 1100 adopts a modular design of columns 1111 and beams 1112. The rigidity of the entire installation constraint frame can be improved by adding diagonal supports, which can constrain and hoist the front and rear mounting sections of the tested aero-engine 3000 onto the entire load-bearing frame. The diagonal supports preferably include an inner triangular support 1114 fixed between the upper end of the column 1111 and the beam 1112, and an outer triangular support 1115 fixed at the lower end of the column 1111.

[0120] like Figure 6 As shown, in a preferred embodiment of the aero-engine whole-machine support stiffness testing device of the present invention, the support connection structure 1220 includes a front support 1230, an inner cylinder 1240, and a rear support 1250 connected in sequence; the support frame 1210 includes a front support frame 1211 and a rear support frame 1212; the front support frame 1211 is connected to the axial end of the front support 1230, and the rear support frame 1212 is connected to the axial end of the rear support 1250.

[0121] like Figure 6As shown, in a preferred embodiment of the aero-engine overall support stiffness testing device of the present invention, the front support 1230 includes an annular first support plate 1231, the inner ring of the first support plate 1231 is connected to one side of the connecting cylinder 1232, the other side of the connecting cylinder 1232 is connected to the outer ring of an annular second support plate 1233, the second support plate 1233 is connected to a plurality of spoke-shaped third support plates 1234, and the plurality of third support plates 1234 are connected to the inner cylinder 1240; the upper end of the front support frame 1211 is provided with a support cylinder 1213, and the front support frame 1211 is connected to the inner ring of the second support plate 1233 through the support cylinder 1213.

[0122] like Figure 6 As shown, in a preferred embodiment of the aero-engine overall support stiffness testing device of the present invention, the rear support 1250 includes a plurality of fourth support plates 1251. One end of the fourth support plate 1251 is vertically connected to the inner cylinder 1240, and the other end of the fourth support plate 1251 is vertically connected to one end of the first connecting plate 1252. The other end of the first connecting plate 1252 is vertically connected to one end of the fifth support plate 1253, and the other end of the fifth support plate 1253 is vertically connected to one end of the second connecting plate 1254. The other end of the second connecting plate 1254 is vertically connected to the outer ring of the annular sixth support plate 1255. A plurality of third connecting plates 1256 are also provided on the outer edge of the outer ring of the sixth support plate 1255. A support cylinder 1213 is provided at the upper end of the rear support frame 1212, and the rear support frame 1212 is connected to the inner ring of the sixth support plate 1255 through the support cylinder 1213.

[0123] The design of the fulcrum displacement measuring device 1200 also incorporates the structural characteristics of the tested aircraft engine 3000.

[0124] Since the first fulcrum 3100 is located inside the fan casing of the tested aero-engine 3000, at the front end of the intermediate casing, its surrounding space is unrestricted, and therefore its displacement can be directly measured. The second fulcrum 3200 is located inside the intermediate casing of the tested aero-engine 3000 and near the front end. Although the space in front is restricted by the bearing housing at the first fulcrum 3100, it can still be measured indirectly by installing a rigid rod. The measurement principle is as follows: Figure 5 As shown, the conversion formulas are as shown in equation (2) and equation (3).

[0125] Z′1=Z1-L2·sinθ (2)

[0126]

[0127] Where Z1 is the vertical displacement at the position directly above the second fulcrum 3200, Z3 is the vertical displacement at the position directly below the second fulcrum 3200, Z′1 and Z′3 are the vertical displacements of the rigid rods extending from positions Z1 and Z3 respectively, X1 and X3 are the horizontal displacements of the rigid rods extending from positions Z1 and Z3 respectively, L2 is the length of the rigid rod, and R2 is half the vertical distance between the rigid rods above and below the second fulcrum 3200.

[0128] The third fulcrum 3300, the fourth fulcrum 3400, and the fifth fulcrum 3500 are all located inside the tested aero-engine 3000, making them difficult to measure using the methods described above. Therefore, a support connection structure 1220 for the fulcrum displacement measuring device 1200 was specially designed. The support connection structure 1220 extends through the entire interior of the tested aero-engine 3000, and is installed simultaneously with the displacement sensor during assembly. The installation design of the support connection structure 1220 is as follows: Figure 6 As shown.

[0129] After the support connection structure 1220 is connected and assembled with the test aircraft engine 3000, the support connection structure 1220 is temporarily fixedly connected to the front end of the fan casing and the rear end of the turbine casing of the test aircraft engine 3000 through the first support plate 1231 and the third connecting plate 1256. This not only positions the support connection structure 1220 and the sensors, but also integrates it with the test aircraft engine 3000, facilitating hoisting and transportation. After the test aircraft engine 3000 is installed on the test stand, support cylinders 1213 and support frames 1210 are installed on its front and rear sides based on the ground. The support frame 1210 is fixedly connected to the front and rear of the support connection structure 1220 through the support cylinders 1213. Subsequently, the first support plate 1231 and the third connecting plate 1256 connecting the front end of the fan casing and the rear end of the turbine casing of the test aircraft engine 3000 are removed, realizing the change of the support method of the support connection structure 1220, and also realizing the installation of the internal pivot displacement sensor of the test aircraft engine 3000.

[0130] like Figure 4 As shown, in a preferred embodiment of the aero-engine whole-engine fulcrum stiffness test device of the present invention, the fulcrum loading fixture 1300 is disposed on one side of the axial direction of the aero-engine under test 3000. The fulcrum loading fixture 1300 includes an actuation component 1310, a loading lever 1320, and a support point component 1330.

[0131] The first end 1321 of the loading lever passes through and is connected to the interior of the test aircraft engine 3000 along the axial direction; the actuation component 1310 is connected to the second end 1322 of the loading lever and is used to apply force to the second end 1322 of the loading lever; the support point component 1330 is connected to the middle of the loading rod.

[0132] Preferably, the first end 1321 of the loading lever is provided with a fulcrum loading disk 1323, and the first end 1321 of the loading lever is connected and fixed inside the test aircraft engine 3000 through the fulcrum loading disk 1323.

[0133] like Figure 4 As shown, in a preferred embodiment of the aero-engine whole-machine support stiffness test device of the present invention, the actuation component 1310 includes an actuation device 1311, a loading end force sensor 1312 and a loading end bearing assembly 1313.

[0134] The actuating device 1311 is connected to the loading end force sensor 1312, the loading end force sensor 1312 is connected to the loading end bearing assembly 1313, and the loading end bearing assembly 1313 is connected to the second end 1322 of the loading lever.

[0135] like Figure 4 As shown, in a preferred embodiment of the aero-engine whole-machine support stiffness testing device of the present invention, the support point assembly 1330 includes a support point pad 1331, a support point bearing assembly 1332, and a support point force sensor 1333.

[0136] The support point pad 1331 is connected to the support point bearing assembly 1332, the support point bearing assembly 1332 is connected to the support point force sensor 1333, and the support point force sensor 1333 is connected to the middle of the loading lever 1320.

[0137] The fulcrum loading fixture 1300 is designed based on the structural characteristics of the fulcrum of the tested aero-engine 3000. The fulcrum loading fixture 1300 adopts lever loading and is a universal loading fixture. It can achieve the loading effect of all fulcrums simply by changing the loading parts.

[0138] like Figure 1 As shown, in a preferred embodiment of the aero-engine overall support stiffness test device of the present invention, a thrust tie rod assembly 3800 is provided between the front mounting section and the rear mounting section of the tested aero-engine 3000.

[0139] The thrust tie rod assembly 3800 is used to connect and fix the front mounting section and the rear mounting section of the test aircraft engine 3000.

[0140] As described above, the aero-engine whole-machine support stiffness testing device of the present invention not only solves the problem that the displacement of the support point inside the aero-engine is difficult to measure directly, but also solves the problem of difficult test support structure design and installation by installing the support connection structure 1220 of the support displacement measuring device 1200 inside the aero-engine under test 3000 and positioning the displacement measuring point and the support connection structure 1220 synchronously. Furthermore, it takes into account the support point of the missile support structure 2100 on the aero-engine whole-machine and supplements the design of the missile support stiffness testing fixture 2000 to measure the missile support stiffness.

[0141] like Figure 8 As shown, the present invention also provides a design method for testing the overall support stiffness of an aero-engine. This design method is used to design the aero-engine overall support stiffness testing device described above. The design method includes the following steps:

[0142] Step S1: Design constraint fixture 1100 based on the actual hoisting state of the tested aero-engine 3000;

[0143] Step S2: Design the pivot loading fixture 1300 based on the structural characteristics and loading requirements of the test aero-engine 3000.

[0144] Step S3: Design a fulcrum displacement measuring device 1200 according to the requirements for fulcrum displacement measurement and the test installation status;

[0145] Step S4: Design a 2000-meter spring support stiffness test fixture;

[0146] Step S5: Construct a test method system for the overall support stiffness of aero-engines.

[0147] It should be noted that the constraint fixture 1100 designed in step S1 can be used to hoist the entire machine onto the modular frame of columns 1111 and beams 1112 via the front and rear mounting sections.

[0148] It should be noted that the fulcrum loading fixture 1300 designed in step S2 can change the direction of the force to achieve load loading by lever loading.

[0149] It should be noted that the spring support stiffness test fixture 2000 designed in step S4 can be used to calculate the accurate stiffness at the rotor support using a formula.

[0150] The aero-engine whole-machine support stiffness test design method of the present invention integrates a whole-machine installation constraint design, loading design, measurement design, data processing and other test method designs, which can comprehensively and systematically realize the stiffness test of each support point of the whole machine.

[0151] The whole - machine support point loading test tooling 1000 designed by the aero - engine whole - machine support point stiffness test design method of the present invention is installed and positioned synchronously during the aero - engine whole - machine assembly process. By changing the support method, the installation of the whole - machine support point loading test tooling 1000 on the test bench is realized. The whole - machine support point loading test tooling 1000 does not interfere with the whole - machine casing, and the installation is flexible.

[0152] The elastic support stiffness test device designed by the aero - engine whole - machine support point stiffness test design method of the present invention realizes the installation constraint and loading of the elastic support structure 2100, measures the elastic support stiffness value, and further obtains the support point stiffness of the elastic support structure 2100 installed on the aero - engine whole - machine.

[0153] The present invention also provides an aero - engine whole - machine support point stiffness test method. The test method uses the aero - engine whole - machine support point stiffness test device as described above for testing. The test method includes the following steps:

[0154] Step P1: The tested aero - engine 3000 is installed and connected to the whole - machine support point loading test tooling 1000.

[0155] Step P2: Conduct the support point displacement measurement test of the whole tested aero - engine 3000, measure the displacements at each support point, record the relationship between the displacements of each support point and the load, and obtain the stiffness of each support point by linearly fitting the data.

[0156] Step P3: Simulate the actual force on the elastic support structure 2100 of the tested aero - engine 3000 through the elastic support stiffness test tooling 2000 and continuously load the elastic support structure 2100, conduct the elastic support stiffness test of the tested aero - engine 3000, record the relationship between the load and the displacement, and obtain the elastic support stiffness by linearly fitting the data.

[0157] It should be noted that when conducting the support point displacement measurement test in Step P2, since the first support point 3100 is located inside the fan casing of the tested aero - engine 3000, at the front end of the intermediate casing, and the surrounding space is not restricted, the displacement of the first support point 3100 can be directly measured. The second support point 3200 is located inside the intermediate casing of the tested aero - engine 3000 and near the front end. Although the front space is restricted by the bearing housing at the first support point 3100, it can still be indirectly measured by installing a rigid rod. The measurement principle is as Figure 5 shown, and the conversion formula is as follows:

[0158] Z′1 = Z1 - L2·sinθ

[0159]

[0160] Where Z1 is the vertical displacement at the position directly above the second fulcrum 3200, Z3 is the vertical displacement at the position directly below the second fulcrum 3200, Z′1 and Z′3 are the vertical displacements of the rigid rods extending from positions Z1 and Z3 respectively, X1 and X3 are the horizontal displacements of the rigid rods extending from positions Z1 and Z3 respectively, L2 is the length of the rigid rod, and R2 is half the vertical distance between the rigid rods above and below the second fulcrum 3200.

[0161] The third fulcrum 3300, the fourth fulcrum 3400, and the fifth fulcrum 3500 are all located inside the tested aero-engine 3000, making them difficult to measure using the methods described above. Therefore, a support connection structure 1220 for the fulcrum displacement measuring device 1200 was specially designed. The support connection structure 1220 extends through the entire interior of the tested aero-engine 3000, and is installed simultaneously with the displacement sensor during assembly. The installation design of the support connection structure 1220 is as follows: Figure 6 As shown.

[0162] After the support connection structure 1220 is connected and assembled with the test aircraft engine 3000, the support connection structure 1220 is temporarily fixedly connected to the front end of the fan casing and the rear end of the turbine casing of the test aircraft engine 3000 through the first support plate 1231 and the third connecting plate 1256. This not only positions the support connection structure 1220 and the sensors, but also integrates it with the test aircraft engine 3000, facilitating hoisting and transportation. After the test aircraft engine 3000 is installed on the test stand, support cylinders 1213 and support frames 1210 are installed on its front and rear sides based on the ground. The support frame 1210 is fixedly connected to the front and rear of the support connection structure 1220 through the support cylinders 1213. Subsequently, the first support plate 1231 and the third connecting plate 1256 connecting the front end of the fan casing and the rear end of the turbine casing of the test aircraft engine 3000 are removed, realizing the change of the support method of the support connection structure 1220, and also realizing the installation of the internal pivot displacement sensor of the test aircraft engine 3000.

[0163] It should be noted that, during the spring support stiffness test in step P3, in order to accurately obtain the stiffness at the rotor support, it is necessary to further obtain the stiffness of the spring support structure 2100, and calculate the complete stiffness at the support point using the following formula:

[0164]

[0165] Wherein, K is the complete stiffness at the fulcrum, K1 is the stiffness from the installation point of the elastic support structure 2100 to the constraint point of the installation section, and K2 is the stiffness of the cascaded elastic support at the fulcrum.

[0166] To measure the stiffness of the projectile support, a projectile support stiffness testing fixture 2000 was designed to simulate the actual force on the projectile support and achieve continuous loading. Figure 7 As shown. Preferably, the spring support structure 2100 is constrained to a rigid mounting plate. The spring support loading disk 2200 simulates the contact between the bearing and the spring support structure 2100 and is axially positioned within the spring support. A first loading rod 2300 is inserted through the center of the spring support loading disk 2200. Connecting pieces 2400 are symmetrically distributed at both ends of the first loading rod 2300 with the axial midpoint of the spring support loading disk 2200 as the center. The connecting pieces 2400 are connected to the upper second loading rod 2500. A loading actuator 2600 is set at the middle position of the upper second loading rod 2500 to achieve continuous loading. Preferably, displacement sensors are arranged as needed on the outer surface of the spring support structure 2100 to record the relationship between the load and each displacement. The spring support stiffness can be obtained by linearly fitting the data.

[0167] The present invention provides a test method for the stiffness of the support points of an aero-engine as a whole. By using a test device for the stiffness of the support points of an aero-engine as a whole, the method solves the problem that the displacement of the internal support points of the aero-engine is difficult to measure directly. Furthermore, it takes into account the support points of the missile support structure 2100 mounted on the aero-engine as a whole and measures the stiffness of the missile support structure 2100.

[0168] The aero-engine whole-machine support stiffness test method of the present invention can effectively obtain the stiffness of all support points of the aero-engine whole machine. It can further verify the conformity between the whole machine casing finite element model and the actual structure through support rotor dynamics analysis, and then verify the rationality of the simplification method of the whole machine casing model, element type selection, mesh size setting and typical connection stiffness and other key modeling parameters, so as to provide experimental support for subsequent engine modeling work.

[0169] In summary, the aero-engine overall support stiffness test design method of the present invention includes a complete set of test method designs such as overall engine installation constraint design, loading design, measurement design, and data processing, which can comprehensively and systematically realize the stiffness test of each support point of the entire engine. The overall engine support loading test fixture 1000 in the aero-engine overall support stiffness test device and test design method of the present invention is installed and positioned synchronously during the overall engine assembly process. The installation of the overall engine support loading test fixture 1000 on the test bench is achieved by changing the support method. The overall engine support loading test fixture 1000 does not interfere with the overall engine casing, and the installation is flexible. The projectile support stiffness test device in the aero-engine overall support stiffness test device and test design method of the present invention realizes the installation constraint and loading of the projectile support structure 2100, can measure the projectile support stiffness value, and thus obtain the support stiffness of the projectile support structure 2100 mounted on the entire engine.

[0170] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. A test device for the overall support stiffness of an aero-engine, characterized in that, The aero-engine whole-engine support stiffness testing device includes a whole-engine support loading test fixture, which includes: a constraint fixture, a support displacement measuring device, and a support loading fixture. The constraint fixture is used to fix the test aircraft engine. The constraint fixture includes a fixing component and a hoisting device. The fixing component is arranged to surround the test aircraft engine, and the hoisting device is connected and fixed to the test aircraft engine. The fulcrum displacement measuring device includes a support frame and a support connection structure. The support connection structure is installed inside the tested aero-engine and provides multiple force application fulcrums for the tested aero-engine. The support frame is used to support the support connection structure. The fulcrum loading fixture is disposed on one side of the axial direction of the tested aero-engine. The fulcrum loading fixture extends into and connects to the interior of the support connection structure along the axial direction of the support connection structure. The fulcrum loading fixture is used to apply a radial force to the support connection structure.

2. The aero-engine overall support stiffness testing device as described in claim 1, characterized in that, The aero-engine whole-machine support stiffness test device also includes a missile support stiffness test fixture, which can simulate the actual force on the missile support and can continuously apply force to the missile support. The spring support stiffness testing fixture includes a spring support structure, a spring support loading plate is provided on the side of the spring support structure, a first loading rod is provided through the center of the spring support structure and the spring support loading plate, the two ends of the first loading rod are connected to one end of a connector, the other end of the connector is connected to the two ends of a second loading rod, and a loading actuator is provided in the middle of the second loading rod.

3. The aero-engine overall support stiffness testing device as described in claim 1, characterized in that, The fixing assembly includes at least two columns and a crossbeam. The hoisting device includes a hanging plate. The two columns are fixedly installed on both radial sides of the test aircraft engine. The two ends of the crossbeam are connected to the upper ends of the two columns. The crossbeam is installed on the upper part of the test aircraft engine. The hanging plate is installed on the lower part of the crossbeam and is connected to the upper part of the test aircraft engine.

4. The aero-engine overall support stiffness testing device as described in claim 3, characterized in that, The columns include at least two front columns and at least two rear columns, the crossbeams include a front crossbeam and a rear crossbeam, and the hanging plates include a front mounting section hanging plate and a rear mounting section hanging plate; The two ends of the front crossbeam are connected to the upper ends of the two front columns, and the two ends of the rear crossbeam are connected to the upper ends of the two rear columns. A longitudinal beam connects the front columns and the rear columns. A front mounting section hanger plate is provided at the lower part of the front crossbeam and is connected to the upper part of the front mounting section of the tested aero-engine. A rear mounting section hanger plate is provided at the lower part of the rear crossbeam and is connected to the upper part of the rear mounting section of the tested aero-engine.

5. The aero-engine overall support stiffness testing device as described in claim 1, characterized in that, The support connection structure includes a front support, an inner cylinder, and a rear support connected in sequence; the support frame includes a front support frame and a rear support frame; the front support frame is connected to the axial end of the front support, and the rear support frame is connected to the axial end of the rear support.

6. The aero-engine overall support stiffness testing device as described in claim 5, characterized in that, The front support includes a first annular support plate, the inner ring of which is connected to one side of a connecting cylinder, and the other side of the connecting cylinder is connected to the outer ring of a second annular support plate. The second support plate is connected to a plurality of spoke-shaped third support plates, which are connected to the inner cylinder. A support cylinder is provided at the upper end of the front support frame, and the front support frame is connected to the inner ring of the second support plate through the support cylinder.

7. The aero-engine overall support stiffness testing device as described in claim 5, characterized in that, The rear support includes multiple fourth support plates. One end of each fourth support plate is vertically connected to the inner cylinder, and the other end of each fourth support plate is vertically connected to one end of a first connecting plate. The other end of the first connecting plate is vertically connected to one end of a fifth support plate, and the other end of the fifth support plate is vertically connected to one end of a second connecting plate. The other end of the second connecting plate is vertically connected to the outer ring of a circular sixth support plate. Multiple third connecting plates are also provided on the outer edge of the outer ring of the sixth support plate. A support cylinder is provided at the upper end of the rear support frame, and the rear support frame is connected to the inner ring of the sixth support plate through the support cylinder.

8. The aero-engine overall support stiffness testing device as described in claim 1, characterized in that, The fulcrum loading fixture is disposed on one side of the axial direction of the tested aero-engine, and the fulcrum loading fixture includes an actuation component, a loading lever, and a support point component; The first end of the loading lever passes axially through and is connected inside the tested aircraft engine; the actuation assembly is connected to the second end of the loading lever and is used to apply force to the second end of the loading lever; the support point assembly is connected to the middle of the loading lever.

9. The aero-engine overall support stiffness testing device as described in claim 8, characterized in that, The actuation assembly includes an actuation device, a loading end force sensor, and a loading end bearing assembly; The actuating device is connected to the loading end force sensor, the loading end force sensor is connected to the loading end bearing assembly, and the loading end bearing assembly is connected to the second end of the loading lever.

10. The aero-engine overall support stiffness testing device as described in claim 8, characterized in that, The support point assembly includes a support point pad, a support point bearing assembly, and a support point force sensor; The support point pad is connected to the support point bearing assembly, the support point bearing assembly is connected to the support point force sensor, and the support point force sensor is connected to the middle of the loading lever.

11. The aero-engine overall support stiffness testing device as described in claim 1, characterized in that, A thrust rod assembly is provided between the front mounting section and the rear mounting section of the tested aero-engine.

12. A method for designing test results for the overall support stiffness of an aero-engine, characterized in that, The design method is used to design the aircraft engine overall support stiffness testing device as described in any one of claims 1-11, and the design method includes the following steps: S1. Design constraint fixtures based on the actual hoisting state of the tested aero-engine; S2. Design the load-bearing fixture for the support point based on the structural characteristics and loading requirements of the entire support point of the tested aero-engine. S3. Design a fulcrum displacement measuring device based on the fulcrum displacement measurement requirements and the test installation status; S4. Design the test fixture for spring support stiffness; S5. Construct a test method system for the overall support stiffness of aero-engines.

13. A method for testing the overall support stiffness of an aero-engine, characterized in that, The test method uses the aircraft engine whole-frame support stiffness test device as described in any one of claims 1-11, and the test method includes the following steps: P1. The tested aircraft engine is installed and connected to the whole aircraft pivot loading test fixture; P2. Conduct a support displacement measurement test on the entire test aircraft engine, measure the displacement at each support point, record the relationship between the displacement at each support point and the load, and obtain the stiffness of each support point by linear fitting the data. P3. Simulate the actual stress on the missile support structure of the tested aero-engine by using a missile support stiffness test fixture and continuously load the missile support structure to conduct a missile support stiffness test on the tested aero-engine. Record the relationship between load and displacement, and obtain the missile support stiffness by linear fitting the data.

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