Aircraft sonic boom experimental model and sonic boom wind tunnel test system and method thereof

Abstract

The invention belongs to the technical field of aircraft acoustic explosion wind tunnel tests, and particularly relates to an aircraft acoustic explosion experiment model and an acoustic explosion wind tunnel test system and method thereof. Comprising a front fuselage model, a middle fuselage model with wings, a fuselage afterbody model with an empennage, a fuselage afterbody model without an empennage, a supporting rod, a blocking block and a horizontal measuring platform. The rear end of the front fuselage model is detachably connected with the front end of the middle fuselage model with the wings; the front end of the fuselage afterbody model with the empennage or the fuselage afterbody model without the empennage is detachably connected with the rear end of the middle fuselage model with the wings; the front end of the supporting rod is detachably connected to the back of the middle fuselage model with the wings, and the rear end is connected to the supporting structure; the back of the rear end of the front fuselage model is provided with a mounting hole, and the plugging block or the horizontal measuring platform is mounted in the mounting hole.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft sonic boom wind tunnel testing, and specifically relates to an aircraft sonic boom experimental model and a sonic boom wind tunnel testing system and method thereof. Background Art

[0002] Supersonic flight can effectively reduce long-distance flight time, improve operational efficiency and passenger comfort, and has gradually become a key direction of civil aircraft development.

[0003] Supersonic flight will produce a huge sonic boom, which is a major problem restricting the development of supersonic aircraft. At present, the main means of studying the sonic boom of supersonic flight of aircraft include numerical simulation, wind tunnel testing, and flight verification. Among them, wind tunnel testing is not only a verification of the correctness of numerical simulation, but also an important support for flight verification research. It plays a vital role in sonic boom research.

[0004] Currently, the models used in aircraft sonic boom wind tunnel tests are difficult to assemble and assemble, difficult to adjust the shape, and not convenient to set up and adjust in the wind tunnel. The test measurement efficiency and accuracy are low, and the test cost is high.

[0005] This application is proposed in view of the above-mentioned technical defects. Summary of the invention

[0006] The purpose of the present application is to provide an aircraft sonic boom test model and a sonic boom wind tunnel test system and method thereof, so as to overcome or alleviate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] An aircraft sonic boom test model comprises a front fuselage model, a middle fuselage model with wings, a rear fuselage model with tail fins, a rear fuselage model without tail fins, a support rod, a block, and a horizontal measurement platform;

[0009] The rear end of the front fuselage model is detachably connected to the front end of the middle fuselage model with wings;

[0010] The front end of the rear fuselage model with a tail or the rear fuselage model without a tail is detachably connected to the rear end of the middle fuselage model with wings;

[0011] The front end of the strut is detachably connected to the back of the middle fuselage model with wings, and the rear end is connected to the supporting structure;

[0012] The rear end of the front fuselage model has a mounting hole, and a block or a horizontal measuring platform is installed in the mounting hole. When installing and debugging the aircraft sonic boom test model, the horizontal measuring platform is installed in the mounting hole to observe the angle of attack of the aircraft sonic boom test model. When conducting sonic boom wind tunnel test measurements, the block is installed in the mounting hole to fill the mounting hole.

[0013] According to at least one embodiment of the present application, in the above-mentioned aircraft sonic boom test model, the rear end of the front fuselage model has a connecting ear piece, and the front end of the middle fuselage model with wings has a connecting ear hole, the connecting ear piece is inserted into the connecting ear hole, and is connected by bolts. The bolts are arranged on the side of the middle fuselage model with wings, and the head is flush with the outer surface of the middle fuselage model with wings.

[0014] According to at least one embodiment of the present application, in the above-mentioned aircraft sonic boom test model, the front ends of the fuselage rear body model with tail wing and the fuselage rear body model without tail wing have connecting columns, and the rear end of the middle fuselage model with wing has connecting holes;

[0015] When the rear fuselage model with tail wing and the middle fuselage model with wing are connected, the connecting column at the rear end of the rear fuselage model with tail wing is inserted into the connecting hole and connected by bolts. The bolts are arranged at the bottom of the middle fuselage model with wing, and the heads are flush with the outer surface of the middle fuselage model with wing.

[0016] When the rear fuselage model without the tail wing is connected to the middle fuselage model with the wing, the connecting column at the rear end of the rear fuselage model without the tail wing is inserted into the connecting hole and connected by bolts. The bolts are arranged at the bottom of the middle fuselage model with the wing, and the heads are flush with the outer surface of the middle fuselage model with the wing.

[0017] According to at least one embodiment of the present application, in the above-mentioned aircraft sonic boom test model, the back of the mid-fuselage model with wings has a strip hole, the front end of the support rod has a strip block, the strip block is arranged in the strip hole, connected by bolts, and two blocking strips are arranged in the strip hole;

[0018] The two blocking strips are fitted with the support rods and connected to the middle fuselage model with wings through bolts, and the two blocking strips and the bolt heads thereof are flush with the outer surface of the middle fuselage model with wings.

[0019] According to at least one embodiment of the present application, in the above-mentioned aircraft sonic boom test model, a force measuring balance may be provided at the rear end of the support rod.

[0020] According to at least one embodiment of the present application, in the above-mentioned aircraft sonic boom test model, the rear end of the strut is tapered, and the angle between its axis and the front end mounting surface is 5 degrees;

[0021] The angle between the front mounting surface of the strut and the horizontal plane of the aircraft sonic boom test model is 2.5 degrees;

[0022] The angle between the horizontal plane of the aircraft sonic boom test model and the rear end axis of the strut is 2.5 degrees.

[0023] A sonic boom wind tunnel test system, comprising an aircraft sonic boom test model, a model motion attitude adjustment mechanism, a sonic boom detection device, and a data acquisition and processing module;

[0024] The sonic explosion detection device is set at the bottom of the wind tunnel;

[0025] The aircraft sonic boom test model is connected to the model motion adjustment mechanism through the rear end of the support rod. The aircraft sonic boom test model is located above the sonic boom detection device, and is 1-3 times the model aircraft length above the sonic boom detection device.

[0026] The model motion adjustment mechanism can drive the aircraft sonic boom test model to move axially, approach or move away from the sonic boom detection device in the axial direction, and adjust the pitch attitude of the aircraft sonic boom test model to change the angle of attack of the aircraft sonic boom test model;

[0027] The data acquisition and processing module is connected to the sonic boom detection device to collect and process the near-field spatial pressure distribution measurement data of the aircraft sonic boom experimental model.

[0028] According to at least one embodiment of the present application, in the above-mentioned sonic explosion wind tunnel test system, the model motion posture adjustment mechanism includes a guide rail, a motion bracket, a pitch adjustment block, and a suspension rod;

[0029] The rails are attached to the bottom of the wind tunnel;

[0030] The moving bracket is connected to the guide rail and can slide axially along the guide rail;

[0031] The pitch adjustment block is hinged on the motion bracket and can rotate;

[0032] One end of the suspension rod is connected to the pitch adjustment block, and the other end is connected to the rear end of the support rod.

[0033] According to at least one embodiment of the present application, in the above-mentioned sonic explosion wind tunnel test system, an actuator is arranged between the moving bracket and the guide rail, and the actuator can drive the moving bracket to slide axially along the guide rail through the extension and retraction of the piston rod, and the control end of the actuator is connected to the wind tunnel measurement and control system, and the extension and retraction of the piston rod is controlled by the wind tunnel measurement and control system. At the same time, the wind tunnel measurement and control system can adjust the wind speed in the wind tunnel.

[0034] According to at least one embodiment of the present application, in the above-mentioned sonic explosion wind tunnel test system, an actuator cylinder is arranged between the pitch adjustment block and the motion bracket, and the actuator cylinder can drive the pitch adjustment block to rotate by the extension and retraction of the piston rod, and the control end of the actuator cylinder is connected to the wind tunnel measurement and control system, and the extension and retraction of the piston rod is controlled by the wind tunnel measurement and control system.

[0035] According to at least one embodiment of the present application, in the above-mentioned sonic explosion wind tunnel test system, the sonic explosion detection device includes a plurality of sound pressure measurement points arranged axially at the front end of the guide rail.

[0036] According to at least one embodiment of the present application, in the above-mentioned sonic explosion wind tunnel test system, the front end of the guide rail can be specifically designed to have a measuring section, the cross-sectional height of the measuring section is 350 mm, the top is rounded, and a total of 404 sound pressure measuring points with an inner diameter of 0.4 mm are arranged along the axis of the top, and the theoretical spacing between adjacent sound pressure measuring points is 4 mm.

[0037] A sonic explosion wind tunnel test method, comprising:

[0038] Carry out data collection for measurement trips, place the aircraft sonic boom experimental model above the sound pressure measurement point, conduct sonic boom wind tunnel tests, and measure the initial spatial pressure distribution data;

[0039] Carry out reference train data collection, move the aircraft sonic boom test model outside the measurement area or dismantle it, and measure the reference space pressure distribution data;

[0040] The spatial pressure change caused by the aircraft sonic boom experimental model is obtained by subtracting the reference spatial pressure distribution data from the initial spatial pressure distribution data, that is, the near-field sonic boom overpressure distribution.

[0041] According to at least one embodiment of the present application, in the above-mentioned sonic boom wind tunnel test method, when carrying out measurement vehicle data collection, the aircraft sonic boom test model is controlled to move axially at a certain interval above the sound pressure measurement point, and N positions are measured. Initial spatial pressure distribution data and reference spatial pressure distribution data are collected once at each position to obtain N groups of spatial pressure changes caused by aircraft sonic boom test models. The spatial pressure changes caused by these N groups of aircraft sonic boom test models are aligned in position and then arithmetic averaged, and the result obtained is the spatial pressure change caused by the aircraft sonic boom test model after spatial averaging. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] Figure 1 is a schematic diagram of an aircraft sonic boom experimental model provided in an embodiment of the present application;

[0043] Figure 2 It is a schematic diagram of the cooperation between the front fuselage model and the middle fuselage model with wings provided in the embodiment of the present application;

[0044] Figure 3 It is a schematic diagram of the coordination of the front fuselage model, the middle fuselage model with wings, and the rear fuselage model with tail provided in the embodiment of the present application;

[0045] Figure 4 It is a schematic diagram of the coordination of the front fuselage model, the middle fuselage model with wings, and the rear fuselage model without tail wing provided in the embodiment of the present application;

[0046] Figure 5 is a schematic diagram of a support rod and a blocking strip thereof provided in an embodiment of the present application;

[0047] Figure 6 is a schematic diagram of the cooperation between the front fuselage model and the horizontal measurement platform provided in an embodiment of the present application;

[0048] Figure 7 It is a schematic diagram of the cooperation between the front fuselage model and the block provided in the embodiment of the present application;

[0049] Figure 8 is a schematic diagram of a sonic explosion wind tunnel test system provided in an embodiment of the present application;

[0050] Fig. 9 It is a schematic diagram of the use of the model motion posture adjustment mechanism provided in the embodiment of the present application;

[0051] Fig.10 is a schematic diagram of a near-field sonic boom signal repeatability test provided in an embodiment of the present application;

[0052] Fig.11 It is a schematic diagram of the comparison between the sonic boom test provided in the embodiment of the present application and the sonic boom signal calculated by CFD;

[0053] Fig.12 is a schematic diagram of flow field characteristics obtained by CFD calculation provided in an embodiment of the present application;

[0054] Fig.13 It is a schematic diagram of the comparison between the sonic boom test provided in the embodiment of the present application and the sonic boom signal calculated by CFD;

[0055] Fig.14 It is a schematic diagram comparing sonic boom signals of the full-machine configuration and the wing-body assembly configuration without a tail provided in the embodiment of the present application;

[0056] Fig.15 1 is a schematic diagram of comparing sonic boom signals of a wing-body assembly configuration test without a tail wing at different angles of attack provided in an embodiment of the present application;

[0057] in:

[0058] 1-front fuselage model; 2-middle fuselage model with wings; 3-rear fuselage model with tail; 4-rear fuselage model without tail; 5-support rod; 6-blocking bar; 7-blocking block; 8-horizontal measuring platform; 9-guide rail; 10-motion bracket; 11-pitch adjustment block; 12-suspension rod; 13-measuring section.

[0059] In order to better illustrate the present embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and should not be construed as limitations on the present application. DETAILED DESCRIPTION

[0060] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be described in further detail in detail and in detail with reference to the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design.

[0061] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall have the usual meanings understood by those of ordinary skill in the art to which this application belongs. The words used in the description of this application to indicate orientation are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. The word "include" used in the description of this application indicates that the elements or objects appearing before the word include the elements or objects listed after the word and their equivalents, but does not exclude other elements or objects.

[0062] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installation", "connection" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Technical personnel in the field can understand its specific meaning in this application according to the specific circumstances.

[0063] An aircraft sonic boom experimental model, such as Figure 1 As shown, it includes a front fuselage model 1, a middle fuselage model 2 with wings, a rear fuselage model 3 with a tail, a rear fuselage model 4 without a tail, a support rod 5, a block 7, and a horizontal measurement platform 8.

[0064] The rear end of the front fuselage model 1 is detachably connected to the front end of the middle fuselage model 2 with wings. Specifically, the rear end of the front fuselage model 1 can be designed to have a connecting ear piece, and the front end of the middle fuselage model 2 with wings can have a connecting ear hole. The connecting ear piece is inserted into the connecting ear hole and connected by bolts. The bolts can be set on the side of the middle fuselage model 2 with wings, and the head is flush with the outer surface of the middle fuselage model 2 with wings, such as Figure 2 shown.

[0065] The front end of the rear fuselage model 3 with a tail or the rear fuselage model 4 without a tail is detachably connected to the rear end of the middle fuselage model 2 with wings. The front end of the rear fuselage model 3 with a tail or the rear fuselage model 4 without a tail can be connected to the rear end of the middle fuselage model 2 with wings according to the specific actual test settings. Specifically, the front ends of the rear fuselage model 3 with a tail and the rear fuselage model 4 without a tail can be designed to have connecting columns, and the rear end of the middle fuselage model 2 with wings can have connecting holes. When the rear fuselage model 3 with a tail is connected to the middle fuselage model 2 with wings, the connecting columns at the rear end of the rear fuselage model 3 with a tail are inserted into the connecting holes and connected by bolts. The bolts can be set at the bottom of the middle fuselage model 2 with wings, and the heads are flush with the outer surface of the middle fuselage model 2 with wings, such as Figure 3 As shown; when the rear fuselage model 4 without the tail wing and the middle fuselage model 2 with the wing are connected, the connecting column at the rear end of the rear fuselage model 4 without the tail wing is inserted into the connecting hole and connected by bolts. The bolts can be set at the bottom of the middle fuselage model 2 with the wing, and the head is flush with the outer surface of the middle fuselage model 2 with the wing, as shown in FIG. Figure 4 shown.

[0066] The front end of the support rod 5 is detachably connected to the back of the middle fuselage model 2 with wings, and the rear end is connected to the supporting structure. Specifically, the back of the middle fuselage model 2 with wings can be designed to have a strip hole, the front end of the support rod 5 can have a strip block, the strip block is set in the strip hole, and connected with bolts, and two blocking strips 6 are set in the strip hole, and the two blocking strips 6 are in contact with the support rod 5, such as Figure 5 As shown, the two blocking strips 6 and the bolt heads thereof are flush with the outer surface of the middle fuselage model 2 with the wing. A force measuring balance can be arranged at the rear end of the support rod 5.

[0067] The rear end of the front fuselage model 1 has a mounting hole, and the block 7 or the horizontal measurement platform 8 is installed in the mounting hole. Specifically, when installing and debugging the aircraft sonic boom test model, the horizontal measurement platform can be installed in the mounting hole to observe the angle of attack of the aircraft sonic boom test model, such as Figure 6 As shown, when performing sonic explosion wind tunnel test measurement, the block 7 is installed in the mounting hole and the mounting hole is filled, as shown in FIG. Figure 7 shown.

[0068] The rear end of the support rod 5 is tapered, and the angle between its axis and the front end mounting surface is 5 degrees. The angle between the front end mounting surface of the support rod 5 and the horizontal plane of the aircraft sonic boom test model is 2.5 degrees. The angle between the horizontal plane of the aircraft sonic boom test model and the rear end axis of the support rod 5 is also 2.5 degrees.

[0069] Based on the aircraft sonic boom test model disclosed in the above embodiment, the present application further provides a sonic boom wind tunnel test system, such as Figure 8As shown, it includes an aircraft sonic boom experimental model, a model motion attitude adjustment mechanism, a sonic boom detection device, and a data acquisition and processing module.

[0070] Sonic boom wind tunnel test is an important means to conduct research on sonic booms of supersonic aircraft. Due to the limitation of wind tunnel size, it is impossible to directly simulate the far field size. It mainly measures the near-field detached pressure distribution. The measured pressure spatial distribution is used as the input condition of the far-field propagation model, and then the far-field sonic boom overpressure is obtained.

[0071] The sonic boom detection device is arranged at the bottom of the wind tunnel, and the aircraft sonic boom test model is connected to the model motion adjustment mechanism through the rear end of the support rod 5. The aircraft sonic boom test model is located above the sonic boom detection device, and is 1-3 times of the model length above the sonic boom detection device. The model motion adjustment mechanism can drive the aircraft sonic boom test model to move axially, approach or move away from the sonic boom detection device in the axial direction, and adjust the pitch attitude of the aircraft sonic boom test model, and change the angle of attack of the aircraft sonic boom test model. The data acquisition and processing module is connected to the sonic boom detection device to collect and process the measurement data of the near-field spatial pressure distribution of the aircraft sonic boom test model.

[0072] The model motion posture adjustment mechanism can be specifically designed to include a guide rail 9, a motion bracket 10, a pitch adjustment block 11, and a suspension rod 12, such as Fig. 9 shown.

[0073] The guide rail 9 is connected to the bottom of the wind tunnel. In order to reduce the impact on the wind tunnel, the guide rail 9 can be designed to have a smaller thickness.

[0074] The motion bracket 10 is connected to the guide rail 9 and can slide axially along the guide rail 9. Specifically, an actuator can be arranged between the motion bracket 10 and the guide rail 9, and the actuator can drive the motion bracket 10 to slide axially along the guide rail 9 through the extension and retraction of the piston rod. Further, the control end of the actuator can be designed to be connected to the wind tunnel measurement and control system, and the extension and retraction of the piston rod can be controlled by the wind tunnel measurement and control system, and the wind tunnel measurement and control system can adjust the wind speed in the wind tunnel.

[0075] The pitch adjustment block 11 is hinged on the motion bracket 10 and can rotate. Specifically, an actuator can be arranged between the pitch adjustment block 11 and the motion bracket 10, and the actuator can drive the pitch adjustment block 11 to rotate by the extension and retraction of the piston rod. Further, the control end of the actuator can be designed to be connected to the wind tunnel measurement and control system, and the extension and retraction of the piston rod can be controlled by the wind tunnel measurement and control system.

[0076] One end of the suspension rod 12 is connected to the pitch adjustment block 11 , and the other end is connected to the rear end of the support rod 5 .

[0077] The moving bracket 10 slides axially along the guide rail 9, and can drive the aircraft sonic boom test model to move axially closer to or farther away from the sonic boom detection device. The pitch adjustment block 11 rotates to adjust the pitch attitude of the aircraft sonic boom test model and change the angle of attack of the aircraft sonic boom test model.

[0078] The sonic explosion detection device includes a plurality of sound pressure measuring points arranged axially at the front end of the guide rail 9. Specifically, the front end of the guide rail 9 can be designed to have a measuring section 13, the section height of the measuring section is 350mm, the top is rounded, and a total of 404 sound pressure measuring points with an inner diameter of 0.4mm are arranged along the axis of the top, and the theoretical spacing between adjacent sound pressure measuring points is 4mm. The pressure measurement adopts the Optimus system, and the module range includes 5psid, 15psid, 45psid and 75psid, and the full-scale accuracy is 0.05%.

[0079] During the sonic boom wind tunnel test, the aircraft sonic boom test model can be controlled to move axially, with a maximum moving distance of 2100mm. Each sonic boom test wind tunnel runs for about 180 seconds, with a total of 20 measurement steps, and the pressure distribution of the model sonic boom influence area at different spatial positions can be measured.

[0080] The sonic explosion wind tunnel test is essentially the measurement of the near-field spatial pressure distribution of the test model, which is generally selected at 1-3 times the length of the model. This type of test has high requirements for the quality of the flow field. In a supersonic wind tunnel, the sonic explosion wind tunnel test of a supersonic aircraft is limited by the space of the wind tunnel test section, the model size is small, and the spatial inhomogeneity and temporal unsteadiness of the wind tunnel flow field have a great impact on the test measurement accuracy. In addition, support interference, model vibration, shock wave reflection, boundary layer interference, etc. are also important factors that affect the measurement accuracy of wind tunnel tests, which need to be fully considered in test design, testing, and data processing.

[0081] Based on the above embodiments of the present application, a sonic explosion wind tunnel test method is further provided, which is as follows:

[0082] Carry out data collection for measurement trips, place the aircraft sonic boom test model above the sound pressure measurement point, conduct sonic boom wind tunnel tests, and measure the initial spatial pressure distribution data of the aircraft sonic boom test model and all components including track 9 in the flow field;

[0083] Carry out reference train data collection, move the aircraft sonic boom experimental model outside the measurement area or dismantle it, and measure the reference space pressure distribution data when only track 9 is in the flow field;

[0084] The spatial pressure change caused by the aircraft sonic boom experimental model is obtained by subtracting the reference spatial pressure distribution data from the initial spatial pressure distribution data, that is, the near-field sonic boom overpressure distribution.

[0085] The reference train method assumes that there is no pressure fluctuation in the wind tunnel itself, and relies on interference such as the buckle track 9 to obtain the net pressure distribution of the aircraft sonic boom test model. Due to the existence of pressure disturbances in the flow field, the effect of the reference train method is not very ideal. In order to reduce the adverse effects of non-uniform disturbances in the wind tunnel flow field on the sonic boom signal measurement, it is also necessary to use spatial averaging technology to further improve the measurement accuracy of the sonic boom test.

[0086] When carrying out data collection for the measurement trip, the aircraft sonic boom experimental model is controlled to move axially at a certain interval above the sound pressure measuring point, and N positions are measured. Initial spatial pressure distribution data and reference spatial pressure distribution data are collected once at each position, and the spatial pressure changes caused by N groups of aircraft sonic boom experimental models are obtained according to the following formulas. The spatial pressure changes caused by these N groups of aircraft sonic boom experimental models are aligned in position and then arithmetic averaged. The result obtained is the spatial pressure change caused by the aircraft sonic boom experimental model after spatial averaging, that is, the near-field pressure distribution after spatial averaging.

[0087]

[0088] in,

[0089] The spatial pressure changes caused by the aircraft sonic boom experimental model;

[0090] It is the overpressure when the aircraft sonic boom test model is located in the measurement area;

[0091] It is the overpressure outside the pressure measurement area of ​​the aircraft sonic boom test model;

[0092] p rail is the pressure measured at the pressure measuring point, p ∞ is the free flow static pressure of the wind tunnel.

[0093] The aircraft sonic boom test model and its sonic boom wind tunnel test system and method disclosed in the above-mentioned embodiments are used to conduct sonic boom tests. The assembly is convenient, the shape can be easily adjusted, the setting and adjustment in the wind tunnel are convenient, the test measurement efficiency and accuracy are high, and the test cost can be greatly reduced.

[0094] In a specific example, in order to verify the stability of the entire test system, a near-field sonic boom signal repeatability test of the whole-machine configuration was carried out on the processed aircraft sonic boom experimental model. The test Mach number was selected to be 1.6, the model angle of attack was 3.5°, and the sonic boom signal was measured at a position 1 times the body length directly below the model. The test results are as follows: Fig.10As shown. The test measurement steps of train No. 1 and train No. 2 are both 8, with a step interval of 4mm. The test data is the result after spatial averaging and zero drift correction. It can be seen that the overall law of the sonic boom signal is basically the same, and the sonic boom signal law generated by the wing part of the two train tests is basically the same. The test data curve has good consistency, and the maximum deviation does not exceed 0.003, indicating that the test stability of the sonic boom measurement system disclosed in this application is good, the test result deviation is small, and it can meet the needs of aircraft sonic boom testing.

[0095] The test Mach number was selected as 1.6, the model angle of attack was 3.5°, and the sonic boom signal was measured at a position 1 times the body length directly below the model. The sonic boom test of the full-machine configuration was carried out and compared with the sonic boom signal calculated by CFD. Fig.11 The wind tunnel test has 8 measurement steps, with a step interval of 4 mm. The data is spatially averaged and zero drift corrected. It can be seen that the overall law of the wind tunnel test of the whole machine configuration is consistent with that of the CFD sonic boom signal. In the interval of x / l=0.9-1.2, the sonic boom signal is in good agreement, including the peak value of the head wave and the law of signal change. This interval corresponds to the clean appearance of the low sonic boom fuselage of the model, without any screw holes and components such as the horizontal measurement platform installation area. In the interval of x / l=1.2-1.41, the sonic boom signal has the same law, but there are certain differences in the value. This interval corresponds to the blade support on the back of the model or the upstream of the wing, including the installation area of ​​the horizontal measurement platform, the front and rear interfaces of the model, and the front section of the modified cover plate on the upper surface of the model that wraps the blade support, which are more or less different from the theoretical shape. In the interval of x / l>1.41, the sonic boom signal is obviously different, especially in terms of the maximum peak value of the signal. The CFD result is about 0.01, while the test result is about 0.018. This interval corresponds to the model wing, blade support, vertical tail and high horizontal tail. The main reason for the difference may be the interference of the back support of the model.

[0096] In order to further analyze the influence of support interference, CFD is used to calculate the near-field sonic boom signal of the aircraft sonic boom experimental model with support, and compared with the wind tunnel test value. The flow field characteristics calculated by CFD are as follows: Fig.12 As shown in the figure, the comparison with the sonic boom test signal is as follows: Fig.13 As shown in the figure, it can be seen from the calculation results that the numerical simulation results with support are in good agreement with the wind tunnel test results. The difference in the peak value of the waveform signal is between 0.003 and 0.004. There is a certain difference between the calculated value of the rear body wave system and the test value, which shows that the support changes the pressure distribution near the wing, thereby increasing the peak value of the near-field overpressure distribution. It is necessary to strengthen the processing of support interference in the test results.

[0097] Based on the aircraft sonic boom experimental model, a sonic boom wind tunnel test was conducted on the wing-body assembly without a tail. The test Mach number was 1.6, the model angle of attack was 3.5°, and the sonic boom signal was measured at a position 1 times the body length directly below the model. The sonic boom signals of the full-machine configuration and the wing-body assembly configuration without a tail were compared, as shown in the figure. Fig.14 As shown, the measurement steps of the wing-body combination wind tunnel test model without tail are 5, and the interval between the steps is 4mm. The data is spatially averaged and zero drift corrected. It can be seen that the sonic boom signal in the range of x / l=0.9-1.77 is generally consistent, with certain differences only in local positions. The change of the sonic boom signal in the range of x / l=1.77-1.9 reflects the influence of the vertical tail and the high horizontal tail on the sonic boom signal. It can be seen from the test data that the absolute value of the negative peak value of the signal in the range of the wing-body combination without tail is larger, that is, the airflow expansion is stronger, which is consistent with the idea that in the design of the rear body sonic boom reduction of the aerodynamic layout of supersonic aircraft, the leading edge shock wave generated by the high horizontal tail is easier to merge with the expansion wave of the wing trailing edge during propagation to reduce the sonic boom intensity. From the perspective of aerodynamic coefficients, the CL corresponding to the full configuration is 0.108, and the CL corresponding to the wing-body combination is 0.104. The lift coefficient corresponding to the wing-body combination configuration without the contribution of the high horizontal tail is slightly smaller.

[0098] On the basis of the sonic boom wind tunnel test of the wing-body assembly without tail, the sonic boom wind tunnel test considering the influence of the angle of attack was carried out by changing the incoming flow angle of attack. The test Mach number was 1.6, the model angle of attack was 3.5° and 4.5°, and the sonic boom signal was measured at a position of 1 times the body length directly below the model. The sonic boom signal comparison curve of the wing-body assembly configuration test without tail is as follows: Fig.15 As shown in the figure, there are 5 measurement steps for the wind tunnel test with a model angle of attack of 4.5°, with a step interval of 4mm. The data is spatially averaged and zero drift corrected. It can be seen that with the increase of the angle of attack, the sonic boom signal shows a "compression" trend as a whole, the sonic boom signal influence area is shortened, and the sonic boom signal generated at the same model position moves upstream. With the increase of the angle of attack, the intensity of the model sonic boom signal increases. The peak value of the sonic boom signal in the range of x / l=1.47-1.48 increases from 0.0169 to 0.0204, with a change of about 0.003. From the perspective of aerodynamic coefficients, the CL corresponding to the model angle of attack of 4.5° is 0.133, and the CL corresponding to the model angle of attack of 3.5° is 0.104. The contribution of the angle of attack to the lift coefficient is obvious.

[0099] Comprehensive test results show that the sonic boom wind tunnel test results can better reflect the distribution law of sonic boom signals of different models such as the whole aircraft and wing-body combination, as well as the influence of different incoming flow conditions such as Mach number and angle of attack. The measurement accuracy of sonic boom signals of supersonic civil aircraft is high and the results are reliable.

[0100] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. An aircraft sonic boom experimental model, characterized in that: It includes a front fuselage model (1), a middle fuselage model with wings (2), a rear fuselage model with tail wing (3), a rear fuselage model without tail wing (4), a support rod (5), a block (7), and a horizontal measurement platform (8); The rear end of the front fuselage model (1) and the front end of the middle fuselage model (2) with wings are connected in a detachable manner; The front end of the rear fuselage model with a tail (3) or the rear fuselage model without a tail (4) is detachably connected to the rear end of the middle fuselage model with wings (2); The front end of the support rod (5) is detachably connected to the back of the middle fuselage model (2) with wings, and the rear end is connected to the supporting structure; The rear end of the front fuselage model (1) is provided with a mounting hole, and a blocking block (7) or a horizontal measurement platform (8) is installed in the mounting hole. When installing and debugging the aircraft sonic boom test model, the horizontal measurement platform is installed in the mounting hole to observe the angle of attack of the aircraft sonic boom test model. When performing sonic boom wind tunnel test measurement, the blocking block (7) is installed in the mounting hole to fill the mounting hole.

2. The aircraft sonic boom experimental model according to claim 1, characterized in that: The front fuselage model (1) has a connecting ear piece at the rear end, and the middle fuselage model with wings (2) has a connecting ear hole at the front end. The connecting ear piece is inserted into the connecting ear hole and connected by bolts. The bolts are arranged on the side of the middle fuselage model with wings (2), and the head is flush with the outer surface of the middle fuselage model with wings (2).

3. The aircraft sonic boom experimental model according to claim 2, characterized in that: The front ends of the rear fuselage model with tail wing (3) and the rear fuselage model without tail wing (4) are provided with connecting columns, and the rear end of the middle fuselage model with wing (2) is provided with connecting holes; When the rear fuselage model with tail wing (3) and the middle fuselage model with wing (2) are connected, the connection column at the rear end of the rear fuselage model with tail wing (3) is inserted into the connection hole and connected by bolts. The bolts are arranged at the bottom of the middle fuselage model with wing (2), and the heads of the bolts are flush with the outer surface of the middle fuselage model with wing (2); When the rear fuselage model (4) without a tail wing is connected to the middle fuselage model (2) with wings, the connection column at the rear end of the rear fuselage model (4) without a tail wing is inserted into the connection hole and connected by bolts. The bolts are arranged at the bottom of the middle fuselage model (2) with wings, and the heads of the bolts are flush with the outer surface of the middle fuselage model (2) with wings.

4. The aircraft sonic boom experimental model according to claim 3, characterized in that: The middle fuselage model (2) with wings has a strip hole on its back, and the front end of the support rod (5) has a strip block, which is arranged in the strip hole and connected with bolts, and two blocking strips (6) are arranged in the strip hole; The two blocking strips (6) are fitted with the support rods (5) and are connected to the middle fuselage model (2) with wings by means of bolts, and the two blocking strips (6) and the bolt heads thereof are flush with the outer surface of the middle fuselage model (2) with wings.

5. The aircraft sonic boom experimental model according to claim 4, characterized in that: A force measuring balance can be arranged at the rear end of the support rod (5).

6. The aircraft sonic boom experimental model according to claim 5, characterized in that: The rear end of the support rod (5) is tapered, and the angle between its axis and the front end mounting surface is 5 degrees; The angle between the front end mounting surface of the support rod (5) and the horizontal plane of the aircraft sonic boom test model is 2.5 degrees; The angle between the horizontal plane of the aircraft sonic boom test model and the rear end axis of the support rod (5) is 2.5 degrees.

7. A sonic explosion wind tunnel test system, characterized in that: It includes an aircraft sonic boom experimental model, a model motion attitude adjustment mechanism, a sonic boom detection device, and a data acquisition and processing module; The sonic explosion detection device is set at the bottom of the wind tunnel; The aircraft sonic boom test model as claimed in the claim is connected to the model motion adjustment mechanism through the rear end of the support rod (5), and the aircraft sonic boom test model is located above the sonic boom detection device and is 1-3 times the model aircraft length above the sonic boom detection device; The model motion adjustment mechanism can drive the aircraft sonic boom test model to move axially, approach or move away from the sonic boom detection device in the axial direction, and adjust the pitch attitude of the aircraft sonic boom test model to change the angle of attack of the aircraft sonic boom test model; The data acquisition and processing module is connected to the sonic boom detection device to collect and process the near-field spatial pressure distribution measurement data of the aircraft sonic boom experimental model.

8. The sonic explosion wind tunnel test system according to claim 7, characterized in that: The model motion posture adjustment mechanism comprises a guide rail (9), a motion bracket (10), a pitch adjustment block (11), and a suspension rod (12); The guide rail (9) is connected to the bottom of the wind tunnel; The moving bracket (10) is connected to the guide rail (9) and can slide axially along the guide rail (9); The pitch adjustment block (11) is hinged on the motion bracket (10) and can rotate; One end of the suspension rod (12) is connected to the pitch adjustment block (11), and the other end is connected to the rear end of the support rod (5).

9. The sonic explosion wind tunnel test system according to claim 8, characterized in that: An actuator is arranged between the moving bracket (10) and the guide rail (9). The actuator can drive the moving bracket (10) to slide axially along the guide rail (9) by extending and retracting the piston rod. The control end of the actuator is connected to the wind tunnel measurement and control system. The extension and retraction of the piston rod is controlled by the wind tunnel measurement and control system. At the same time, the wind tunnel measurement and control system can adjust the wind speed in the wind tunnel.

10. The sonic explosion wind tunnel test system according to claim 9, characterized in that: An actuator is arranged between the pitch adjustment block (11) and the motion bracket (10), and the actuator can drive the pitch adjustment block (11) to rotate by the extension and retraction of a piston rod. The control end of the actuator is connected to a wind tunnel measurement and control system, and the extension and retraction of the piston rod is controlled by the wind tunnel measurement and control system.

11. The sonic explosion wind tunnel test system according to claim 10, characterized in that: The sonic explosion detection device comprises a plurality of sound pressure measuring points arranged axially at the front end of a guide rail (9).

12. The sonic explosion wind tunnel test system according to claim 11, characterized in that: Specifically, the guide rail (9) can be designed to have a measuring section (13) at the front end, the measuring section has a cross-sectional height of 350 mm, a rounded top, and a total of 404 sound pressure measuring points with an inner diameter of 0.4 mm are arranged along the axis of the top, and the theoretical spacing between adjacent sound pressure measuring points is 4 mm.

13. A sonic explosion wind tunnel test method, implemented based on the sonic explosion wind tunnel test system according to claim 12, characterized in that: include: Carry out data collection for measurement trips, place the aircraft sonic boom experimental model above the sound pressure measurement point, conduct sonic boom wind tunnel tests, and measure the initial spatial pressure distribution data; Carry out reference train data collection, move the aircraft sonic boom test model outside the measurement area or dismantle it, and measure the reference space pressure distribution data; The spatial pressure change caused by the aircraft sonic boom experimental model is obtained by subtracting the reference spatial pressure distribution data from the initial spatial pressure distribution data, that is, the near-field sonic boom overpressure distribution.

14. The sonic explosion wind tunnel test method according to claim 13, characterized in that: When carrying out data collection for the measurement trip, the aircraft sonic boom experimental model is controlled to move axially at a certain interval above the sound pressure measuring point, and N positions are measured. Initial spatial pressure distribution data and reference spatial pressure distribution data are collected once at each position to obtain N groups of spatial pressure changes caused by aircraft sonic boom experimental models. The spatial pressure changes caused by these N groups of aircraft sonic boom experimental models are aligned in position and then arithmetic averaged. The result obtained is the spatial pressure change caused by the aircraft sonic boom experimental model after spatial averaging.

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