A low-level lightning test method for aircraft engine systems

Abstract

The present invention relates to the technical field of aircraft lightning testing, and discloses a whole-machine lightning low-level test method applicable to an aircraft engine system, the method comprising the steps of keeping the aircraft engine electrically insulated from the ground through an insulating support frame, constructing a coaxial return conductor loop around the aircraft engine, wherein the front end of the conductor loop is connected to a pulse current generator through a metal lead, and the rear end of the conductor loop is connected to a terminal matching impedance through a metal lead. The present invention simulates a lightning discharge waveform through a pulse current generator, combines a coaxial return conductor loop and a terminal matching impedance, verifies and assesses the lightning protection of an aircraft engine and its internal electrical and electronic systems and equipment, clarifies test requirements, test procedures, test waveforms, etc., forms a unified and executable whole-machine lightning low-level test method for an aircraft engine system, and reduces the cost of conducting whole-machine lightning tests on the engine.

Description

Technical Field

[0001] The invention relates to the technical field of aircraft lightning testing, and more specifically to a whole-aircraft lightning low-level testing method applicable to an aircraft engine system. Background Art

[0002] The aircraft engine system is one of the most critical systems on the aircraft. If a fault that is difficult to recover from occurs during flight, such as a mid-air shutdown, it will have catastrophic consequences for the flight safety of the aircraft. Therefore, whether for fixed-wing aircraft or rotary-wing aircraft, the area where the engine is located is generally regarded as Zone 1, which has the most stringent lightning protection environment, that is, the area with the highest lightning protection requirements. Therefore, in the airworthiness certification of civil aircraft, lightning protection for engine systems has become a mandatory clause, such as CCAR-33.

[0003] With the continuous development and innovation of aviation engine technology, on the one hand, traditional mechanical control systems have been largely replaced by highly integrated electronic control systems, which has greatly increased the sensitivity of the engine's electronic and electrical systems; on the other hand, composite materials with poor electromagnetic shielding performance are widely used, resulting in a more severe electromagnetic environment inside the engine, and the lightning electromagnetic environment poses an increasingly severe threat to the safety of civil aircraft engine systems.

[0004] At present, conventional engine system full-machine lightning tests basically rely on full-machine lightning tests of the aircraft in full state, which has the disadvantages of long cycle, high cost, and complex test. At the same time, the test is easily restricted by other conditions of the aircraft. The existing lightning test standards (such as SAE ARP 5416A, RTCA / DO-160G, etc.) only include lightning indirect effect test methods at the system level and equipment level of the engine, and lack lightning indirect effect test methods at the full-machine level of the aircraft engine. No relevant patents have been found in China so far. Therefore, it is necessary to establish a full-machine lightning low-level test method suitable for aircraft engine systems, which is used to carry out full-machine lightning indirect effect tests of aircraft engine systems, and provide test means to determine whether the aircraft engine system meets the requirements of lightning protection safety indicators. Summary of the invention

[0005] The present invention provides a whole-machine lightning low-level test method applicable to an aircraft engine system, and aims to solve the following problem: conventional whole-machine lightning test of an engine system basically relies on whole-machine lightning test of the aircraft in a full state, which has the disadvantages of long cycle, high cost, and complex test, and the conduct of the test is easily restricted by other conditions of the aircraft.

[0006] To achieve the above object, the present invention provides the following technical solution: a whole-machine lightning low-level test method applicable to an aircraft engine system, comprising the following steps:

[0007] Step 1: Keep the aircraft engine electrically insulated from the ground;

[0008] Step 2: Ensure the continuity of the main electrical circuit of the aircraft engine and the lightning current path of the rest of the aircraft engine system;

[0009] Step 3: construct a coaxial return conductor loop;

[0010] Step 4: Install the measuring probe on the aircraft engine;

[0011] Step 5: Make electrical connections between the engine lightning current injection point and the engine lightning current outflow point;

[0012] Step 6: ground the generator ground terminal, and connect the pulse current generator output terminal to the front end of the conductor loop;

[0013] Step 7: Send the lightning current in step 6 to the aircraft engine to obtain the test waveform generated by the pulse current generator; if the test waveform meets the requirements, proceed to step 8; if the test waveform does not meet the requirements, adjust the parameters of the pulse current generator until the test waveform meets the requirements;

[0014] Step 8: Charge and discharge the pulse current generator and record the environmental noise waveform;

[0015] Step 9: Start the test, restart the pulse current generator, and obtain and save the test data through the acquisition equipment until the test data required for the engine system test is collected.

[0016] In a preferred embodiment, the test waveform generated by the pulse current generator includes a scaled lightning current component A and a scaled current component H, and the current peak range used is between 1 and 20 kA; in step 2, the aircraft engine is not powered on; in step 3, a coaxial return conductor loop is constructed around the aircraft engine, and the front end of the conductor loop of the coaxial return conductor loop is connected to the pulse current generator through a metal lead, and the rear end of the conductor loop of the coaxial return conductor loop is connected to the terminal matching impedance through a metal lead; in step 5, the engine lightning current injection point is connected to the front end of the conductor loop, and the engine lightning current outflow point is connected to the rear end of the conductor loop; in step 6, the pulse current generator is set on the ground at a safe distance from the aircraft engine, and the output end of the pulse current generator is connected to the front end of the conductor loop through a metal lead; the test waveform in step 7 is obtained by measuring equipment; in step 8, the ambient noise waveform is recorded by a measuring probe, and the measuring probe is placed near the measuring point and connected to the measuring equipment.

[0017] In a preferred embodiment, the test data includes harness induced current, single conductor induced voltage / current, voltage / current noise, injected current, magnetic field coupling voltage, structural voltage and structural surface current, and the environmental noise waveform is less than 10% of the measured induced voltage data.

[0018] In a preferred embodiment, the metal lead is a low-impedance copper flat braided tape, and the measuring probe includes a magnetic field probe for detecting lightning on the surface of an aircraft engine, a current probe for detecting wiring harness induced current, a current probe for detecting single conductor induced current and current noise, and a voltage probe for detecting single conductor induced voltage and voltage noise.

[0019] A test device based on a whole-aircraft lightning low-level test method comprises an insulating support frame, a coaxial return conductor loop and an aircraft engine, wherein the aircraft engine is connected to the insulating support frame through an insulating collar and an insulating sling structure, the insulating collar is fixedly mounted on the outside of the aircraft engine, a plurality of insulating sling structures are fixedly connected to the insulating collar, and the insulating sling structure is fixedly connected to the insulating support frame;

[0020] The coaxial return conductor loop includes a plurality of parallel return conductors, which are parallel and equidistantly distributed, and are arranged parallel to the axis direction of the aircraft engine, the ends of the plurality of parallel return conductors are connected by an annular connecting conductor, and the ends of each parallel return conductor are connected to a connecting conductor, and the connecting conductors at both ends of the parallel return conductors are respectively connected to each other to form a conductor loop front end and a conductor loop rear end;

[0021] End fixing frames are fixedly installed at both ends of the insulating support frame, and the ends of the parallel return conductors are fixedly connected with insulating pull ropes. The insulating pull ropes pass through the end fixing frames and slide with the end fixing frames. The end of the insulating pull rope away from the parallel return conductor is fixedly connected with a counterweight structure.

[0022] In a preferred embodiment, a limiting frame is provided on the outside of the counterweight structure, the limiting frame is provided on a grounding plate and fixedly connected to the ground, the counterweight structure is slidably provided in the limiting frame, a reverse damping member is provided between the counterweight structure and the limiting frame, the reverse damping member includes a trapezoidal block and a damping member, the damping member is fixedly mounted on the limiting frame, a trapezoidal groove is provided at a position of the counterweight structure corresponding to the limiting frame, the trapezoidal block is slidably mounted in the trapezoidal groove, and an inclined mating surface is formed between the side of the trapezoidal block facing away from the damping member and the trapezoidal groove, the distance from the inclined mating surface to the damping member is gradually increased from bottom to top, and a damping structure is provided on the mating surface between the trapezoidal block and the damping member.

[0023] In a preferred embodiment, air ducts are fixedly installed on the end fixing frames at both ends of the insulating support frame, the axis of the air duct is set corresponding to the axis of the aircraft engine, and a blowing fan is set in the air duct at one end corresponding to the air inlet of the aircraft engine, and an exhaust fan is set in the other air duct, and the blowing fan and the exhaust fan are both set away from the aircraft engine, and an atomizing nozzle is also set in the air duct corresponding to the air inlet of the aircraft engine.

[0024] In a preferred embodiment, a gap is formed between the insulating collar and the surface of the aircraft engine, a plurality of insulating spherical parts are arranged on the inner side of the insulating collar, a spherical mounting cavity is arranged inside the insulating collar, and the insulating spherical parts are installed in the spherical mounting cavity. The insulating collar on the aircraft engine is provided with at least two groups, and the two groups of insulating collars are respectively arranged at tapered positions at both ends of the aircraft engine, and the insulating sling structures on the two groups of insulating collars are inclined toward the center of the aircraft engine.

[0025] In a preferred embodiment, a baffle is fixedly mounted on the insulating sling structure, a blocking cover is provided on the baffle, a venting cavity is provided in the baffle at a position corresponding to the insulating sling structure, a constriction is provided at one end of the venting cavity facing the aircraft engine, a uniform gap is formed between the constriction and the insulating sling structure, an air pipe is provided inside the insulating sling structure, a connecting air hole is provided at the position of the air pipe corresponding to the venting cavity, and the air pipe is connected to a high-pressure air pump.

[0026] In a preferred embodiment, an air duct is provided inside the insulating collar, the air duct is connected to the air pipe, an annular cavity is provided near the inner wall of the insulating collar in the spherical mounting cavity, the annular cavity is connected to the air duct through an air guide hole, and a gap is formed between the annular cavity near the aircraft engine and the insulating spherical part.

[0027] The beneficial effects of the present invention are as follows: the present invention simulates lightning discharge waveforms through a pulse current generator, combines a coaxial return conductor loop and terminal matching impedance, verifies and assesses the lightning protection of aircraft engines and their internal electrical and electronic systems and equipment, clarifies test requirements, test procedures, test waveforms and other contents, and forms a unified and executable full-machine lightning low-level test method for aircraft engine systems. On the one hand, a lightning test environment suitable for simulating actual lightning strikes on aircraft engine systems is constructed, thereby reducing the cost of conducting full-machine lightning tests on engines; on the other hand, the present invention effectively solves the problem of the lack of standard method references for full-machine lightning tests on aircraft engine systems, and can provide technical support and reference for full-machine lightning tests on other major aircraft structures (such as aircraft cabin sections, etc.). BRIEF DESCRIPTION OF THE DRAWINGS

[0028] Figure 1 The figure is a schematic diagram of the arrangement of the whole aircraft engine system lightning low-level test of the present invention.

[0029] Figure 2 The present invention is a flow chart of the test method.

[0030] Figure 3 It is a simplified diagram of the coaxial return conductor loop of the present invention.

[0031] Figure 4 This is the experimental waveform of the scaled current component A of the pulse current generator discharge of the present invention.

[0032] Figure 5 This is the experimental waveform of the reduced current component H of the pulse current generator discharge of the present invention.

[0033] Figure 6 The present invention is a flow chart of the testing method for verifying the coaxial return conductor loop structure arrangement.

[0034] Figure 7 It is a schematic diagram of the actual structure of the test device of the present invention.

[0035] Figure 8 The schematic diagram of the test device after adding airflow simulation to the present invention.

[0036] Fig. 9 It is a side view of the counterweight structure of the present invention.

[0037] Fig.10 For the present invention Fig. 9 A-section structure enlarged view.

[0038] Fig.11 It is a side view of the aircraft engine in the test device of the present invention.

[0039] Fig.12 It is an enlarged view of the local structure of the insulating collar of the present invention.

[0040] Fig.13 Another distribution schematic diagram of the insulating sling structure of the present invention.

[0041] Fig.14 The present invention is based on Fig.12 An enlarged view of the structure of the collar on the surface of the aircraft engine.

[0042] Fig.15 The figure is a schematic diagram of the structure after a barrier is added to the insulating sling structure of the present invention.

[0043] Fig.16 It is an enlarged view of the structure of the barrier member of the present invention.

[0044] Fig.17 This is a state diagram of the air flow blowing out around the insulating spherical member of the present invention.

[0045] In the figure: 1. Pulse current generator; 2. Coaxial return conductor loop; 201. Parallel return conductor; 202. Ring connecting conductor; 203. Connecting conductor; 3. Aircraft engine; 4. Output end of pulse current generator; 5. Generator grounding end; 6. Front end of conductor loop; 7. Back end of conductor loop; 8. Engine lightning current injection point; 9. Engine lightning current outflow point; 10. Terminal matching impedance; 11. Fiber optic transceiver; 12. Measuring equipment; 13. Ground; 14. Grounding plate; 15. Insulating support frame; 16. Measuring probe; 17. Low-impedance copper flat braid ; 18. Signal transmission line; 19. Shielding box; 20. Insulating collar; 21. Insulating sling structure; 22. End fixing frame; 23. Insulating pull rope; 24. Insulating spherical part; 25. Spherical installation cavity; 251. Annular cavity; 26. Airway; 27. Air pipe; 28. Air guide hole; 29. ​​Air duct; 30. Blowing fan; 31. Exhaust fan; 32. Atomizing nozzle; 33. Counterweight structure; 34. Limiting frame; 35. Reverse damping part; 351. Trapezoidal block; 352. Damping part; 353. Trapezoidal groove; 36. Partition member; 37. Blocking cover; 38. Air release cavity; 39. Shrinkage. DETAILED DESCRIPTION

[0046] The subject matter described herein will now be discussed with reference to example embodiments. It should be understood that the discussion of these embodiments is only to enable those skilled in the art to better understand and implement the subject matter described herein, and the functions and arrangements of the elements discussed may be changed without departing from the scope of protection of the contents of this specification. Each example may omit, replace or add various processes or components as needed. In addition, the features described relative to some examples may also be combined in other examples.

[0047] In order to evaluate and verify the lightning indirect effect protection capability of the aircraft engine system in the laboratory, it is necessary to Figure 1 The following configurations are carried out in a laboratory or an open field test site:

[0048] (1) The aircraft engine 3 is electrically insulated from the ground 13 through an insulating support frame 15. The insulating support frame 15 used in the example is preferably made of wood. In order to better construct the coaxial return conductor loop 2, the height between the aircraft engine 3 and the ground 13 is 1.2 m.

[0049] (2) Construct a simulated part or structure electrically connected to the aircraft engine 3 and use conductive materials to electrically connect them to simulate the real lightning current path of the aircraft engine 3. The electronic and electrical systems and equipment on the aircraft engine 3 do not need to be powered on, but the continuity of the main electrical circuit needs to be ensured. In this example, a low-impedance copper flat braided belt 17 is used as the electrical connection between the aircraft engine 3 and the aircraft fuselage.

[0050] (3) Follow the instructions Figure 3 As shown, a coaxial return conductor loop 2 is constructed around the aircraft engine 3. The existence of the coaxial return conductor loop 2 can regulate the current distribution on the engine surface to make it consistent with the actual lightning strike of the aircraft engine. In this example, a metal aluminum structure can be used to evenly arrange 16 parallel return conductors 201 at a height of 0.5m outside the aircraft engine. The coaxial return conductor loop 2 uses the same metal aluminum as the annular connecting conductor 202 and the connecting conductor 203 at both ends of the coaxial return conductor loop 2 to form a coaxial return conductor loop 2 together with the parallel return conductor 201.

[0051] (4) A low-impedance copper flat braided tape 17 is used to connect the front end 6 of the return conductor loop to the output end 4 of the pulse current generator.

[0052] (5) A low-impedance copper flat braided tape 17 is used to connect the rear end 7 of the return conductor loop to the terminal matching impedance 10. The terminal matching impedance 10 used in this example is mainly used to eliminate the transmission line reflection effect (mainly caused by the unevenness and discontinuity of the loop formed by the aircraft engine 3 and the coaxial return conductor loop 2), so as to achieve that the test waveform of the pulse current generator 1 acting on the surface of the aircraft engine 3 is not distorted.

[0053] (6) The engine lightning current injection point 8 and the engine lightning current outflow point 9 are selected to simulate the interaction between lightning and the aircraft engine 3, so that the current distribution on the surface of the aircraft engine 3 to be tested can simulate the lightning current distribution when it is struck by lightning in actual flight as much as possible. In this example, one of the lightning strike modes is selected, that is, the engine lightning current injection point 8 is the air intake of the aircraft engine 3, and the engine lightning current outflow point 9 is the tail nozzle of the aircraft engine 3. Therefore, a low-impedance copper flat braided belt 17 is used to connect the engine lightning current injection point 8 to the front end 6 of the return conductor loop, and the engine lightning current outflow point 9 to the rear end 7 of the return conductor loop.

[0054] (7) The pulse current generator 1 is placed on the ground 13 at a safe distance from the aircraft engine 3 (the safe distance refers to the distance at which the pulse current generator 1 and the aircraft engine 3 under test do not affect each other during the test), and the generator ground terminal 5 is connected to the grounding plate 14. The pulse current generator output terminal 4 is connected to the front end 6 of the return conductor loop through a low-impedance copper flat braided belt 17.

[0055] (8) According to the test measurement requirements, the required measuring probes 16 are arranged on the upper part of the aircraft engine 3, and the measuring equipment 12 is arranged. In this embodiment, the data such as the surface lightning of the aircraft engine 3, the wiring harness induced current, the single conductor induced voltage / current, and the voltage / current noise are mainly measured. Among them, the detection of the surface lightning of the aircraft engine 3 mainly adopts the magnetic field probe, the detection of the wiring harness induced current mainly adopts the current probe, and the detection of the single conductor induced voltage / current and the detection of the voltage / current noise mainly adopt the current probe and the voltage probe.

[0056] (9) The lightning current generated by the pulse current generator 1 is sent to the lightning current injection point 8 of the aircraft engine through a low-impedance copper flat braided belt 17. The test waveform generated by the pulse current generator 1 is obtained through the measuring device 12 to confirm whether it meets the requirements. If the test waveform generated by the pulse current generator 1 meets the requirements, the test is carried out; if the test waveform generated by the pulse current generator 1 does not meet the requirements, the parameters of the pulse current generator 1 are adjusted until the test waveform meets the requirements.

[0057] (10) Before the formal test begins, in order to avoid the influence of environmental noise on the test data or the evaluation of the test results, it is necessary to measure the environmental noise waveform of the measured part. In this example, the measuring probe 16 (the measuring probe is selected according to the test data) is placed near the measurement part, the test data is transmitted to the measuring device 12 through the optical fiber transceiver 11, and the aircraft engine 3 is discharged through the pulse current generator 1 to obtain and store the environmental noise waveform.

[0058] (11) The low-level lightning test of the aircraft engine 3 mainly evaluates or verifies the impact of the lightning electromagnetic environment generated when lightning attaches to the surface of the aircraft engine 3 on the electronic and electrical system-level equipment of the aircraft engine. In combination with the lightning protection requirements of the aircraft engine, the test waveforms used in this test include the scaled current component A (scaling factor is 1 / 100) and the scaled current component H (scaling factor is 1 / 10), that is, the current amplitudes of the injected scaled current component A and the scaled current component H are 2kA and 1kA respectively, ensuring that the test current peak range is between 1 and 20kA.

[0059] (12) Complete the measurement of all test data according to the test requirements. In order to avoid the influence of the lightning electromagnetic environment on the test signal transmission line 18 and the measuring equipment 12, this example uses optical fiber or shielded coaxial cable as the signal transmission line 18 during the test, and puts the measuring equipment 12 into the shielding box 19.

[0060] (13) The test data needs to be linearly extrapolated to the actual lightning environment to obtain the actual transient induced levels (ATLs) and transient induced waveforms in the electrical and electronic system cables of the aircraft engine 3, and compared with the transient control levels (TCLs) and transient design levels (ETDLs) of the electrical and electronic systems / equipment, so as to assess or verify whether the lightning protection performance of the aircraft engine 3 meets the requirements.

[0061] After completing the above steps, refer to the instructions attached. Figure 2 The present invention provides a whole-machine lightning low-level test method applicable to an aircraft engine system, which specifically comprises the following steps:

[0062] Step 1: The aircraft engine 3 is electrically insulated from the ground 13 via the insulating support frame 15;

[0063] Step 2: The aircraft engine 3 is not powered on, and the continuity of the main electrical circuit of the aircraft engine 3 and the lightning current path of the rest of the aircraft engine system, such as the pylon, the support rod and the aircraft fuselage, is ensured;

[0064] Step 3: construct a coaxial return conductor loop 2 around the aircraft engine 3, wherein the conductor loop front end 6 of the coaxial return conductor loop 2 is connected to the pulse current generator 1 through a metal lead, and the conductor loop rear end 7 of the coaxial return conductor loop 2 is connected to the terminal matching impedance 10 through a metal lead;

[0065] Step 4: Install the measurement probes 16 required for the test, such as current probes, electric field probes, magnetic field probes, etc., on the aircraft engine 3;

[0066] Step 5: According to the lightning current path of the aircraft engine 3, the positions of the engine lightning current injection point 8 and the engine lightning current outflow point 9 on the aircraft engine 3 are electrically connected to form a lightning current conduction path, wherein the engine lightning current injection point 8 is connected to the front end 6 of the conductor loop, and the engine lightning current outflow point 9 is connected to the rear end 7 of the conductor loop;

[0067] Step 6: The pulse current generator 1 is placed on the ground 13 at a safe distance from the aircraft engine 3, and the generator grounding terminal 5 of the pulse current generator 1 is grounded, and the pulse current generator output terminal 4 of the pulse current generator 1 is connected to the conductor loop front end 6 through a metal lead;

[0068] Step 7, turn on the measuring device 12, send the lightning current in step 6 to the aircraft engine 3 through the metal lead, obtain the test waveform generated by the pulse current generator 1 through the measuring device 12, and confirm whether it meets the requirements; if the test waveform generated by the pulse current generator 1 meets the requirements, proceed to step 8; if the test waveform generated by the pulse current generator 1 does not meet the requirements, adjust the parameters of the pulse current generator 1 until the test waveform meets the requirements;

[0069] Step 8: Place the measuring probe 16 near the measuring point and connect it to the measuring device 12 to charge and discharge the pulse current generator 1 and record the environmental noise waveform;

[0070] Step nine, start the test and restart the pulse current generator 1 to generate a lightning current waveform that meets the test requirements, and obtain and save the test data through the acquisition equipment until the test data required for the engine system to be tested is acquired.

[0071] It should be noted that the accuracy of the above-mentioned measuring probe 16 (such as current probe, electric field probe, magnetic field probe, etc.) is ±5%; the number of parallel return conductors 201 in the coaxial return conductor loop 2 and the height from the surface of the aircraft engine 3 should be determined according to the test requirements. In this embodiment, only 16 parallel return conductors 201 with a height of 0.5m from the surface of the aircraft engine 3 are used as an example for testing. The test waveform generated by the pulse current generator 1 includes a scaled lightning current component A and a scaled current component H. The current peak range used is between 1 and 20kA. Under the premise of not damaging the aircraft engine 3 or equipment, a higher amplitude is used as much as possible. The test data includes harness induced current, single conductor induced voltage / current, voltage / current noise, injection current, magnetic field coupling voltage, structural voltage, structural surface current, etc., wherein the environmental noise waveform should be less than 10% of the measured induced voltage data, but a value up to 20% is acceptable. The harness induced current, single conductor induced voltage / current, voltage / current noise, etc. in the test data need to be linearly extrapolated to the required full amplitude.

[0072] In the above steps, when the scaled current component A and the scaled current component H are used as test waveforms for waveform calibration, the scaled current component A waveform only needs to correct the waveform amplitude, and the scaled current component H waveform needs to comprehensively consider the amplitude and the rate of change of the maximum rising edge.

[0073] In the above steps, the current distribution on the surface of the aircraft engine 3 adopts a coaxial return conductor loop 2, but due to the limitations of test costs and test technology, it is impractical to optimize the structural design of the coaxial return conductor loop 2 solely based on the measured surface current data of the aircraft engine 3. Therefore, in order to effectively and quickly solve the above technical problems, the coaxial return conductor loop 2 in the example of the present invention can achieve structural layout optimization by combining electromagnetic simulation with experimental verification;

[0074] Based on electromagnetic simulation, the influence of the number of parallel return conductors 201 and the height between the parallel return conductors 201 and the surface of the aircraft engine 3 on the current distribution on the surface of the aircraft engine 3 is obtained, so as to theoretically ensure that the current distribution on the surface of the aircraft engine 3 is close to the actual lightning strike situation, and provide an important basis for the arrangement of the coaxial return conductor loop 2 of the low-level lightning test of the entire aircraft engine 3;

[0075] According to the electromagnetic simulation results, a coaxial return conductor loop 2 that meets the test requirements is constructed. The surface current of the external structure of the aircraft engine 3 is obtained through the test to verify whether the structural arrangement of the coaxial return conductor loop 2 is reasonable. For details, refer to the attached manual. Figure 6 , the test method is as follows:

[0076] S1. Prepare a measuring probe 16 and a measuring device 12 (such as a digital oscilloscope, etc.) for measuring the surface current of the structure;

[0077] S2, according to the surface current measurement point specified on the surface of the aircraft engine 3, the measuring probe 16 is installed at the position on the surface of the aircraft engine 3 to be measured, and connected to the measuring device 12 through the signal transmission line 18 and the optical fiber transceiver 11;

[0078] S3. Check whether the test circuit and electrical connection are correct;

[0079] S4, calibration test waveform;

[0080] S5, discharge the pulse current generator 1, record and save the test data, and take test images or photos;

[0081] S6, move the measuring probe 16 to the next measuring position, and repeat steps S4 to S5 until all measurements are completed;

[0082] S7. Based on the measured test data, obtain the current distribution on the surface of the aircraft engine 3, compare it with the electromagnetic simulation data, and comprehensively evaluate whether the structural arrangement of the coaxial return conductor loop 2 meets the test requirements.

[0083] In addition, based on the above test method, the present invention also provides a whole aircraft lightning low level test device suitable for aircraft engine system. Figures 7 to 17, including an insulating support frame 15, a coaxial return conductor loop 2 and an aircraft engine 3, wherein the insulating support frame 15 is arranged on a grounding plate 14, and the specific equipment required for the test and the corresponding connection have been described in the above test method, therefore, this embodiment is only explained with the specific structure of the coaxial return conductor loop 2 and the insulating support frame 15, and the fixation of the aircraft engine 3 as the center, and because in the traditional method, the aircraft engine 3 is mostly placed directly on a simple frame structure, the frame structure forms a barrier under the aircraft engine 3, and due to the existence of the frame under the aircraft engine 3, the air environment above and below the aircraft engine 3 is greatly different, therefore, in In this embodiment, the aircraft engine 3 is connected to the insulating support frame 15 through the insulating collar 20 and the insulating sling structure 21, that is, the insulating support frame 15 is a hanger structure, the insulating collar 20 is fixedly installed on the outside of the aircraft engine 3, and a plurality of insulating sling structures 21 are fixedly connected to the insulating collar 20. One end of the insulating sling structure 21 away from the insulating collar 20 is fixedly connected to the insulating support frame 15, thereby forming a suspended support for the aircraft engine 3, thereby making the environment of the space around the aircraft engine 3 the same and reducing differences, and the insulating sling structures 21 can be evenly distributed, while ensuring that sufficient supporting force is provided, the influence of the insulating sling structures 21 on the space environment is minimized;

[0084] Further, the coaxial return conductor loop 2 includes a plurality of parallel return conductors 201, specifically, 16 of them can be arranged, the plurality of parallel return conductors 201 are parallel and equidistantly distributed, and the parallel return conductors 201 are arranged parallel to the axial direction of the aircraft engine 3, the ends of the plurality of parallel return conductors 201 are connected by an annular connecting conductor 202, and the ends of each parallel return conductor 201 are connected to a connecting conductor 203, and the connecting conductors 203 at both ends of the parallel return conductor 201 are respectively connected to each other to form a conductor loop front end 6 and a conductor loop rear end 7;

[0085] In the above-mentioned embodiment, since the test device is based on the physical structure of the aircraft engine 3 for testing, the structure of the coaxial return conductor loop 2 to be constructed is relatively large, wherein the parallel return conductor 201 has a large span length, and the parallel return conductor 201 itself is easily bent downward due to the influence of gravity, thereby affecting the parallelism of the parallel return conductor 201 and affecting the actual test results. However, directly setting an auxiliary support system outside each parallel return conductor 201 will affect the test environment and the test results. For this reason, refer to the attached manual. Figure 8This embodiment also provides the following technical solutions. Specifically, both ends of the insulating support frame 15 are fixedly installed with end fixing frames 22, and the ends of the parallel return conductors 201 are fixedly connected with insulating pull ropes 23. The insulating pull ropes 23 pass through the end fixing frames 22 and slide with the end fixing frames 22. One end of the insulating pull rope 23 away from the parallel return conductor 201 is fixedly connected with a counterweight structure 33. With the help of the counterweight structure 33, a strong pulling force is provided, so that both ends of the parallel return conductor 201 are subjected to tension, thereby enabling the parallel return conductors 201 to remain in a straight state to avoid bending and affecting the test.

[0086] It should be noted that the insulating collar 20, insulating sling structure 21, end fixing frame 22 and insulating pull rope 23 used in this embodiment are all insulating components. For example, the insulating sling structure 21 and the insulating pull rope 23 can use aramid fiber insulating rope, polyamide insulating rope and glass fiber insulating rope, etc., while the insulating collar 20 and the end fixing frame 22 can use wood or high-strength composite materials (such as carbon fiber reinforced resin composite materials) to avoid their influence on the electric field.

[0087] Furthermore, in the actual test process, the parallel return conductor 201 will form a current. During the test, it is easy for some reasons to cause the current of the parallel return conductor 201 to be too large, which in turn causes the temperature of the parallel return conductor 201 to rise. At this time, if the traction equipment remains unchanged, the parallel return conductor 201 will expand due to heat or the parallel return conductor 201 will become soft, resulting in an increase in length and sagging. Therefore, the gravity vertical pulling scheme of the counterweight structure 33 provided in this embodiment can effectively solve the above problem. However, this requires that the counterweight structure 33 has the ability to always sag and provide elongation compensation for the parallel return conductor 201. The counterweight structure 33 needs to have the ability to move. However, when the counterweight structure 33 is in an active state, the coaxial return conductor loop 2 as a whole is also prone to movement due to some reasons. Therefore, this embodiment also provides the following technical solutions. For details, refer to the attached manual. Fig. 9 and Fig.10A limiting frame 34 is arranged outside the counterweight structure 33, and the limiting frame 34 is arranged on the grounding plate 14 and fixedly connected to the ground 13. The counterweight structure 33 is slidably arranged in the limiting frame 34, and a reverse damping member 35 is arranged between the counterweight structure 33 and the limiting frame 34. The reverse damping member 35 includes a trapezoidal block 351 and a damping member 352, and the damping member 352 is fixedly installed on the limiting frame 34. A trapezoidal groove 353 is arranged at a position of the counterweight structure 33 corresponding to the limiting frame 34, and the trapezoidal block 351 is slidably installed in the trapezoidal groove 353, and an inclined matching surface is formed between the side of the trapezoidal block 351 away from the damping member 352 and the trapezoidal groove 353, and the distance from the inclined matching surface to the damping member 352 is gradually increased from bottom to top, and the trapezoidal block 351 is arranged to be fixedly installed on the limiting frame 34. The mating surface between the trapezoidal block 351 and the damping member 352 is provided with a damping structure, such as a rubber plate. Therefore, when the counterweight structure 33 moves downward, the trapezoidal block 351 can move upward first, and under the action of the inclined mating surface, it is relatively loose with the damping member 352, so that the counterweight structure 33 can descend normally. If the counterweight structure 33 has a tendency to move upward, due to the action of the inclined mating surface, the trapezoidal block 351 will be squeezed toward the damping member 352, thereby increasing the damping of the trapezoidal block 351 and the damping member 352, forming a reverse lock, thereby ensuring that the counterweight structure 33 can automatically move downward to compensate for the unexpected extension of the parallel return conductor 201, and at the same time, the coaxial return conductor loop 2 can be relatively fixed as a whole without causing left and right displacement.

[0088] Furthermore, since the aircraft is relatively moving during navigation, there is relatively flowing airflow around the aircraft engine 3. In order to improve the authenticity of the test environment, this embodiment also simulates the air flow environment. For details, refer to the attached manual. Figure 8 , air ducts 29 are fixedly installed on the end fixing frames 22 at both ends of the insulating support frame 15, and the air ducts 29 are also made of insulating materials. The axis of the air duct 29 corresponds to the axis of the aircraft engine 3, and a blowing fan 30 is arranged in the air duct 29 corresponding to one end of the air inlet of the aircraft engine 3, and an exhaust fan 31 is arranged in the other air duct 29, and the blowing fan 30 and the exhaust fan 31 are both arranged away from the aircraft engine 3. The purpose of setting the air duct 29 is to make the two electrical components of the blowing fan 30 and the exhaust fan 31 away from the test site to reduce the impact, and the blowing fan 30 blows out the airflow, and the exhaust fan 31 draws away the airflow, which can form air flow in the test site, that is, around the aircraft engine 3, to simulate the environment. At the same time, the flow of the above-mentioned airflow can also take away the heat on the surface of the parallel return conductor 201, and timely dissipate the heat of the parallel return conductor 201 to reduce the occurrence of elongation and relaxation of the parallel return conductor 201.

[0089] Furthermore, since the humidity of the surrounding air is relatively high when the aircraft encounters bad weather during navigation, in order to further fit the actual scenario, an atomizing nozzle 32 is also provided in the air duct 29 corresponding to the air inlet of the aircraft engine 3. The atomizing nozzle 32 is used to spray water mist to increase the humidity of the air flowing through the aircraft engine 3.

[0090] In the above embodiment, since it is necessary to form flowing air on the surface of the aircraft engine 3, in order to reduce the influence of the insulating collar 20 on the aircraft engine 3 and reduce the influence of the insulating collar 20 on the electric field on the surface of the aircraft engine 3, it is necessary to reduce the actual contact area between the insulating collar 20 and the aircraft engine 3. For this purpose, refer to the attached manual. Figures 11 to 14 The insulating collar 20 forms a gap with the surface of the aircraft engine 3. A plurality of insulating spherical members 24 are arranged on the inner side of the insulating collar 20. The insulating spherical members 24 are made of an insulating flexible material, such as rubber, or a composite structure in which a hard plastic structure is arranged inside and rubber is arranged outside. A spherical mounting cavity 25 is arranged inside the insulating collar 20. The insulating spherical members 24 are installed in the spherical mounting cavity 25. The insulating spherical members 24 can produce corresponding rolling. When the aircraft engine 3 is fixed, the aircraft engine 3 can be rotated and adjusted to facilitate testing of the aircraft engine 3 in different postures. The insulating collar 20 on the aircraft engine 3 is arranged in at least two groups. Since the front and rear sections of the aircraft engine 3 have a certain taper, the two groups of insulating collars 20 can be arranged respectively on the aircraft engine 3. There are tapered positions at both ends of the aircraft engine 3, and the insulating sling structures 21 on the two groups of insulating collar parts 20 are inclined toward the center of the aircraft engine 3, so that while the insulating sling structure 21 provides support for the aircraft engine 3, the insulating collar part 20 can also be close to the conical surface of the aircraft engine 3, thereby squeezing the insulating spherical part 24. Since the insulating spherical part 24 has a certain flexibility, the friction between it and the spherical mounting cavity 25 can be increased after being squeezed, thereby forming fixation. Since the contact area between the insulating spherical part 24 and the aircraft engine 3 is small, the coverage area of ​​the surface of the aircraft engine 3 can be effectively reduced, thereby reducing the impact on the electric field on the surface of the aircraft engine 3, and at the same time, the impact on the flowing air on the surface of the aircraft engine 3 can also be reduced.

[0091] In the above embodiment, if a high humidity environment simulation is adopted, the air flowing through the outside of the aircraft engine 3 will contain a high moisture content, and the presence of the insulating sling structure 21 will cause part of the moisture to adhere to the surface of the aircraft engine 3. If the moisture flows or distributes greatly, it is easy to form a conductive connection between the insulating sling structure 21 and the insulating support frame 15 or other components. Therefore, in order to reduce the adhesion of water on the surface of the insulating sling structure 21, this embodiment also provides the following technical solutions. For details, refer to the attached manual. Fig.15 and Fig.16A baffle 36 is fixedly installed on the insulating sling structure 21, and a blocking cover 37 is arranged on the baffle 36. A deflation cavity 38 is arranged in the baffle 36 at a position corresponding to the insulating sling structure 21, and a constriction 39 is arranged at one end of the deflation cavity 38 facing the aircraft engine 3, and a uniform gap is formed between the constriction 39 and the insulating sling structure 21. An air pipe 27 is arranged inside the insulating sling structure 21 (when the insulating sling structure 21 is actually manufactured and woven, the air pipe 27 is used as the inner core for weaving, and an insulating composite material is formed outside the air pipe 27 to form the insulating sling structure 21), and a connecting air hole is arranged at the position of the air pipe 27 corresponding to the deflation cavity 38. The air pipe 27 is connected to a high-pressure air pump, and then in actual use, high-pressure gas can be input into the deflation cavity 38 with the help of the air pipe 27, so that the constriction 39 blows out a high-pressure airflow along the surface of the insulating sling structure 21, so as to prevent water vapor from adhering to and flowing on the insulating sling structure 21, thereby improving the safety of the insulating sling structure 21.

[0092] It should be noted that the air pipe 27 may not be set in the insulating sling structure 21, but can be directly connected separately to supply air to the deflation chamber 38, but this solution will occupy external space, and the air pipe 27 is set in the insulating sling structure 21, the gas input by the high-pressure air pump can be high-temperature gas, and the air pipe 27 runs through the insulating sling structure 21, which can heat the insulating sling structure 21 with mechanical energy to keep the insulating sling structure 21 dry.

[0093] Furthermore, due to the high humidity in the air during the test, liquid water is easily gathered around the insulating spherical member 24, thereby affecting the electric field. Fig.12 and Fig.17 An air passage 26 is provided inside the insulating collar 20, and the air passage 26 is communicated with the air pipe 27. An annular cavity 251 is provided near the inner wall of the insulating collar 20 of the spherical mounting cavity 25, and the annular cavity 251 is communicated with the air passage 26 through the air guide hole 28. A gap is formed between the annular cavity 251 and the insulating spherical component 24 near the aircraft engine 3. In actual use, the annular cavity 251 can also blow air outward along the gap around the insulating spherical component 24 to avoid liquid water from gathering around the insulating spherical component 24, thereby reducing the impact on the electric field on the surface of the aircraft engine 3.

[0094] The above-mentioned embodiments only express several implementation methods of the present invention, and the description thereof is relatively specific and detailed, but it cannot be understood as limiting the scope of the present invention. It should be pointed out that, for ordinary technicians in this field, several variations and improvements can be made without departing from the concept of the present invention, which all belong to the protection scope of the present invention.

[0095] The above describes an embodiment of the present embodiment, but the present embodiment is not limited to the above-mentioned specific implementation methods. The above-mentioned specific implementation methods are merely illustrative and not restrictive. Under the guidance of the present embodiment, ordinary technicians in this field can also make many forms, all of which are within the protection of the present embodiment.

Claims

1. A whole-machine lightning low-level test method applicable to an aircraft engine system, characterized in that: The following steps are involved: Step 1: electrically insulate the aircraft engine (3) from the ground (13); Step 2: Ensure the continuity of the main electrical circuit of the aircraft engine (3) and the lightning current path of the rest of the aircraft engine system; Step 3: construct a coaxial return conductor loop (2); Step 4: Installing a measuring probe (16) on the aircraft engine (3); Step 5: electrically connect the engine lightning current injection point (8) and the engine lightning current outflow point (9); Step 6: ground the generator ground terminal (5) and connect the pulse current generator output terminal (4) to the front end of the conductor loop (6); Step 7, sending the lightning current in step 6 to the aircraft engine (3), and obtaining the test waveform generated by the pulse current generator (1); if the test waveform meets the requirements, proceed to step 8; if the test waveform does not meet the requirements, adjusting the parameters of the pulse current generator (1) until the test waveform meets the requirements; Step 8: charging and discharging the pulse current generator (1) and recording the environmental noise waveform; Step 9: Start the test, restart the pulse current generator (1), and acquire and save the test data through the acquisition device until the test data required for the engine system to be tested is acquired; In the second step, the aircraft engine (3) is not powered; in the third step, the coaxial return conductor loop (2) is constructed around the aircraft engine (3), and the conductor loop front end (6) of the coaxial return conductor loop (2) is connected to the pulse current generator (1) through a metal lead, and the conductor loop rear end (7) of the coaxial return conductor loop (2) is connected to the terminal matching impedance (10) through a metal lead; in the fifth step, the engine lightning current injection point (8) is connected to the conductor loop front end (6), and the engine lightning current outflow point (9) is connected to the conductor loop rear end (7); in the sixth step, the pulse current generator (1) is set on the ground (13) at a safe distance from the aircraft engine (3), and the pulse current generator output end (4) is connected to the conductor loop front end (6) through a metal lead.

2. A whole-machine lightning low-level test method applicable to an aircraft engine system according to claim 1, characterized in that: The test waveform generated by the pulse current generator (1) includes a scaled lightning current component A and a scaled current component H, and the current peak value used is in the range of 1 to 20 kA; the test waveform in step seven is obtained by a measuring device (12); in step eight, the environmental noise waveform is recorded by a measuring probe (16), and the measuring probe (16) is placed near a measuring point and connected to the measuring device (12).

3. A whole-machine lightning low-level test method applicable to an aircraft engine system according to claim 2, characterized in that: The test data includes harness induced current, single conductor induced voltage / current, voltage / current noise, injected current, magnetic field coupling voltage, structural voltage and structural surface current. The environmental noise waveform is less than 10% of the measured induced voltage data.

4. A whole-machine lightning low-level test method applicable to an aircraft engine system according to claim 3, characterized in that: The metal lead is a low-impedance copper flat braided tape (17), and the measuring probe (16) comprises a magnetic field probe for detecting lightning on the surface of an aircraft engine (3), a current probe for detecting harness induced current, a current probe for detecting single conductor induced current and current noise, and a voltage probe for detecting single conductor induced voltage and voltage noise.

5. A test device based on the whole machine lightning low level test method according to claim 1, characterized in that: The invention comprises an insulating support frame (15), a coaxial return conductor loop (2) and an aircraft engine (3), wherein the aircraft engine (3) is connected to the insulating support frame (15) via an insulating collar (20) and an insulating sling structure (21), wherein the insulating collar (20) is fixedly mounted on the outside of the aircraft engine (3), a plurality of insulating sling structures (21) are fixedly connected to the insulating collar (20), and the insulating sling structure (21) is fixedly connected to the insulating support frame (15); The coaxial return conductor loop (2) comprises a plurality of parallel return conductors (201), the plurality of parallel return conductors (201) are distributed in parallel and equidistantly, and the parallel return conductors (201) are arranged parallel to the axial direction of the aircraft engine (3), the ends of the plurality of parallel return conductors (201) are connected via an annular connecting conductor (202), and the ends of each parallel return conductor (201) are connected to a connecting conductor (203), and the connecting conductors (203) at both ends of the parallel return conductor (201) are respectively connected to each other to form a conductor loop front end (6) and a conductor loop rear end (7); End fixing frames (22) are fixedly mounted on both ends of the insulating support frame (15); an insulating pull rope (23) is fixedly connected to the end of the parallel return conductor (201); the insulating pull rope (23) passes through the end fixing frame (22) and is slidably engaged with the end fixing frame (22); and a counterweight structure (33) is fixedly connected to one end of the insulating pull rope (23) away from the parallel return conductor (201).

6. A test device based on the whole machine lightning low level test method according to claim 5, characterized in that: A limiting frame (34) is disposed outside the counterweight structure (33); the limiting frame (34) is disposed on a grounding plate (14) and fixedly connected to the ground (13); the counterweight structure (33) is slidably disposed in the limiting frame (34); a reverse damping member (35) is disposed between the counterweight structure (33) and the limiting frame (34); the reverse damping member (35) comprises a trapezoidal block (351) and a damping member (352); the damping member (352) is fixedly mounted on the limiting frame (34). A trapezoidal groove (353) is provided at a position of the counterweight structure (33) corresponding to the limiting frame (34); the trapezoidal block (351) is slidably mounted in the trapezoidal groove (353); and an inclined mating surface is formed between the side of the trapezoidal block (351) facing away from the damping member (352) and the trapezoidal groove (353); the distance from the inclined mating surface to the damping member (352) is gradually increased from bottom to top, and a damping structure is provided on the mating surface between the trapezoidal block (351) and the damping member (352).

7. A test device based on the whole machine lightning low level test method according to claim 6, characterized in that: An air duct (29) is fixedly mounted on the end fixing frames (22) at both ends of the insulating support frame (15); the axis of the air duct (29) is arranged to correspond to the axis of the aircraft engine (3); a blowing fan (30) is arranged in the air duct (29) at one end corresponding to the air inlet of the aircraft engine (3); an exhaust fan (31) is arranged in the other air duct (29); the blowing fan (30) and the exhaust fan (31) are both arranged away from the aircraft engine (3); and an atomizing nozzle (32) is also arranged in the air duct (29) corresponding to the air inlet of the aircraft engine (3).

8. A test device based on the whole machine lightning low level test method according to claim 7, characterized in that: A gap is formed between the insulating collar (20) and the surface of the aircraft engine (3); a plurality of insulating spherical members (24) are arranged on the inner side of the insulating collar (20); a spherical mounting cavity (25) is arranged inside the insulating collar (20); the insulating spherical member (24) is mounted in the spherical mounting cavity (25); the insulating collar (20) on the aircraft engine (3) is provided in at least two groups; the two groups of insulating collars (20) are respectively arranged at tapered positions at both ends of the aircraft engine (3); and the insulating sling structures (21) on the two groups of insulating collars (20) are both arranged to be inclined toward the center of the aircraft engine (3).

9. A test device based on the whole machine lightning low level test method according to claim 8, characterized in that: A baffle (36) is fixedly mounted on the insulating sling structure (21), a blocking cover (37) is arranged on the baffle (36), a deflation cavity (38) is arranged in the baffle (36) at a position corresponding to the insulating sling structure (21), a constriction (39) is arranged at one end of the deflation cavity (38) facing the aircraft engine (3), a uniform gap is formed between the constriction (39) and the insulating sling structure (21), an air pipe (27) is arranged inside the insulating sling structure (21), a connecting air hole is arranged in the air pipe (27) at a position corresponding to the deflation cavity (38), and the air pipe (27) is connected to a high-pressure air pump.

10. A test device based on the whole machine lightning low level test method according to claim 9, characterized in that: An air passage (26) is provided inside the insulating collar (20), the air passage (26) being in communication with an air pipe (27); an annular cavity (251) is provided at a position of the spherical mounting cavity (25) close to the inner wall of the insulating collar (20); the annular cavity (251) is in communication with the air passage (26) via an air guide hole (28); and a gap is formed between the position of the annular cavity (251) close to the aircraft engine (3) and the insulating spherical component (24).

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