An aircraft sonic boom experimental model and a sonic boom wind tunnel test system and method thereof

Abstract

The application belongs to the technical field of aircraft sonic boom wind tunnel test, and particularly relates to an aircraft sonic boom experimental model, a sonic boom wind tunnel test system and a method thereof, wherein the aircraft sonic boom experimental model comprises a front fuselage model, a middle fuselage model with wings, a tail fuselage model with tail wings, a tail fuselage model without tail wings, a support rod, a blocking block and a horizontal measurement platform; the rear end of the front fuselage model is detachably connected with the front end of the middle fuselage model with wings; the front end of the tail fuselage model with tail wings or the tail fuselage model without tail wings is detachably connected with the rear end of the middle fuselage model with wings; the front end of the support rod is detachably connected with the back of the middle fuselage model with wings, and the rear end is connected with a support structure; the rear end of the front fuselage model is provided with a mounting hole, and the blocking block or the horizontal measurement platform is mounted in the mounting hole.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aircraft sonic boom wind tunnel test, and particularly relates to an aircraft sonic boom experimental model and a sonic boom wind tunnel test system and method thereof. BACKGROUND

[0002] Supersonic flight can effectively reduce long-range flight time, improve operation efficiency and passenger comfort, and gradually become the focus of civil aircraft development.

[0003] Supersonic flight produces a great sonic boom, which is an important problem restricting the development of supersonic aircraft. At present, the main means for studying the sonic boom of supersonic aircraft flight are numerical simulation, wind tunnel test, and flight verification. Among them, wind tunnel test is not only a verification of the correctness of numerical simulation, but also an important support for flight verification research, and plays a crucial role in sonic boom research.

[0004] At present, the model used in aircraft sonic boom wind tunnel test is difficult to assemble and adjust the shape, and is not convenient to set and adjust in the wind tunnel, the test measurement efficiency and precision are low, and the test cost is high.

[0005] The present application is proposed in view of the above technical defects. SUMMARY

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

[0007] The technical solution of the present application is:

[0008] An aircraft sonic boom experimental model, comprising a front fuselage model, a middle fuselage model with wings, a fuselage afterbody model with tail wings, a fuselage afterbody model without tail wings, a support rod, a blocking block, and a horizontal measurement platform.

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

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

[0011] The front end of the support rod is detachably connected to the back of the middle fuselage model with wings, and the rear end is connected to a support structure.

[0012] The rear end of the front fuselage model has a mounting hole, and a plug or a horizontal measuring platform is mounted in the mounting hole. During installation and debugging of the aircraft sonic boom experimental model, the horizontal measuring platform is mounted in the mounting hole to observe the angle of attack of the aircraft sonic boom experimental model. During sonic boom wind tunnel test measurement, the plug is mounted in the mounting hole to fill the mounting hole.

[0013] According to at least one embodiment of the present application, the aircraft sonic boom experimental model has a connecting lug at the rear end of the front fuselage model, and a connecting lug hole at the front end of the middle fuselage model with wings. The connecting lug is inserted into the connecting lug hole and connected by a bolt. The bolt is 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, the aircraft sonic boom experimental model has a connecting column at the front end of the tailless fuselage afterbody model and the tail-equipped fuselage afterbody model, and a connecting hole at the rear end of the middle fuselage model with wings.

[0015] When the tail-equipped fuselage afterbody model is connected with the middle fuselage model with wings, the connecting column at the rear end of the tail-equipped fuselage afterbody model is inserted into the connecting hole and connected by a bolt. The bolt is arranged on the bottom of the middle fuselage model with wings, and the head is flush with the outer surface of the middle fuselage model with wings.

[0016] When the tailless fuselage afterbody model is connected with the middle fuselage model with wings, the connecting column at the rear end of the tailless fuselage afterbody model is inserted into the connecting hole and connected by a bolt. The bolt is arranged on the bottom of the middle fuselage model with wings, and the head is flush with the outer surface of the middle fuselage model with wings.

[0017] According to at least one embodiment of the present application, the aircraft sonic boom experimental model has a strip hole at the back of the middle fuselage model with wings, and a strip block at the front end of the support rod. The strip block is arranged in the strip hole and connected by a bolt. Two blocking strips are arranged in the strip hole.

[0018] The two blocking strips are attached to the support rod and connected to the middle fuselage model with wings by the bolt. The two blocking strips and the bolt heads are flush with the outer surface of the middle fuselage model with wings.

[0019] According to at least one embodiment of the present application, the aircraft sonic boom experimental model has a force balance at the rear end of the support rod.

[0020] According to at least one embodiment of the present application, the rear end of the support rod is conical, and the angle between its axis and the mounting surface at the front end is 5 degrees.

[0021] The angle between the mounting surface at the front end of the support rod and the horizontal surface of the aircraft sonic boom experimental model is 2.5 degrees.

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

[0023] The sonic boom wind tunnel test system comprises an aircraft sonic boom experimental model, a model motion posture adjusting mechanism, a sonic boom detection device, and a data acquisition and processing module.

[0024] The sonic boom detection device is arranged at the bottom of the wind tunnel.

[0025] The aircraft sonic boom experimental model is connected to the model motion adjusting mechanism through the rear end of the support rod, and the aircraft sonic boom experimental model is above the sonic boom detection device, at a distance of 1-3 times the model length from the top of the sonic boom detection device.

[0026] The model motion adjusting mechanism can drive the aircraft sonic boom experimental 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 experimental model to change the angle of attack of the aircraft sonic boom experimental model.

[0027] 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 experimental model.

[0028] According to at least one embodiment of the present application, the model motion posture adjusting mechanism in the above-mentioned sonic boom wind tunnel test system comprises a guide rail, a motion support, a pitch adjusting block, and a suspension rod.

[0029] The guide rail is connected to the bottom of the wind tunnel.

[0030] The motion support is connected to the guide rail and can slide axially along the guide rail.

[0031] The pitch adjusting block is hinged to the motion support and can rotate.

[0032] One end of the suspension rod is connected to the pitch adjusting 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, the actuator cylinder is arranged between the motion support and the guide rail, which can drive the motion support to slide axially along the guide rail through 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. At the same time, the wind tunnel measurement and control system can adjust the size of the wind speed in the wind tunnel.

[0034] According to at least one embodiment of the present application, the actuator cylinder is arranged between the pitch adjusting block and the motion support, which can drive the pitch adjusting block to rotate through 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 of the embodiments of the present application, in the above-mentioned sound blast wind tunnel test system, the sound blast detection device comprises a plurality of sound pressure measuring points arranged axially at the front end of the guide rail.

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

[0037] A sound blast wind tunnel test method, comprising:

[0038] Developing measurement train data collection, placing the aircraft sound blast experimental model above the sound pressure measuring points, conducting sound blast wind tunnel test, and measuring initial spatial pressure distribution data;

[0039] Developing reference train data collection, moving the aircraft sound blast experimental model outside the measurement area or removing it, and measuring reference spatial pressure distribution data;

[0040] Subtracting the reference spatial pressure distribution data from the initial spatial pressure distribution data to obtain the spatial pressure change caused by the aircraft sound blast experimental model, i.e. the near-field sound blast overpressure distribution.

[0041] According to at least one of the embodiments of the present application, in the above-mentioned sound blast wind tunnel test method, when developing measurement train data collection, the aircraft sound blast experimental model is controlled to move above the sound pressure measuring points at a certain interval distance along the axial direction, and N positions are measured, and at each position, the initial spatial pressure distribution data and the reference spatial pressure distribution data are collected once, N groups of spatial pressure changes caused by the aircraft sound blast experimental model are obtained, and after position alignment, the N groups of spatial pressure changes caused by the aircraft sound blast experimental model are arithmetically averaged to obtain the spatial pressure change caused by the aircraft sound blast experimental model after spatial averaging. BRIEF DESCRIPTION OF DRAWINGS

[0042] Figure 1 is a schematic diagram of an aircraft sound blast experimental model provided by an embodiment of the present application;

[0043] Figure 2 is a schematic diagram of a front fuselage model cooperating with a middle fuselage model with wings provided by an embodiment of the present application;

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

[0045] Figure 4 is a schematic diagram of a front fuselage model, a middle fuselage model with wings, and a fuselage rear body model without tail provided by an embodiment of the present application;

[0046] Figure 5 is a schematic diagram of a support rod and its plug provided by an embodiment of the present application;

[0047] Figure 6 is a schematic diagram of a front fuselage model cooperating with a horizontal measurement platform provided by an embodiment of the present application;

[0048] Figure 7 is a schematic diagram of a front fuselage model cooperating with a plug provided by an embodiment of the present application;

[0049] Figure 8 is a schematic diagram of a sound blast wind tunnel test system provided by an embodiment of the present application;

[0050] Figure 9 is a schematic diagram of use of a model motion posture adjusting mechanism provided by an embodiment of the present application;

[0051] Figure 10 is a schematic diagram of a near-field sound blast signal repeatability test provided by an embodiment of the present application;

[0052] Figure 11 is a schematic diagram of a sound blast test test, and comparison of sound blast signals of CFD calculation provided by an embodiment of the present application;

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

[0054] Figure 13 is a schematic diagram of a sound blast test test, and comparison of sound blast signals of CFD calculation provided by an embodiment of the present application;

[0055] Figure 14 is a schematic diagram of comparison of test sound blast signals of a full-mechanism type and a wing-body combination configuration without tail provided by an embodiment of the present application;

[0056] Figure 15 is a schematic diagram of comparison of test sound blast signals of a wing-body combination configuration without tail at different angles of attack provided by an embodiment of the present application;

[0057] wherein:

[0058] 1-front fuselage model; 2-middle fuselage model with wings; 3-fuselage rear body model with tail; 4-fuselage rear body model without tail; 5-support rod; 6-plug; 7-plug; 8-horizontal measurement platform; 9-guide rail; 10-motion support; 11-pitch adjusting block; 12-hanging rod; 13-measurement section.

[0059] In order to better illustrate the embodiments, some components in the drawings can be omitted, enlarged or reduced, and do not represent the actual product size. In addition, the drawings are only used for illustrative description and cannot be understood as a limitation of the present application. DETAILED DESCRIPTION

[0060] In order to make the technical solutions of the present application and the advantages thereof clearer, the technical solutions of the present application will be further clearly and completely described in detail below in conjunction with the drawings. It should be understood that the specific embodiments described herein are only partial embodiments of the present application, and are only used to explain the present application, but not to limit the present application. It should be noted that, for the convenience of description, only parts related to the present application are shown in the drawings, and other related parts can be referred to the general design.

[0061] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of the present application should be the general meanings understood by the general technical personnel in the field of the present application. The words indicating the direction used in the description of the present application are only used to indicate the relative direction or positional relationship, and when the absolute position of the described object changes, the relative positional relationship may also change accordingly. The "comprising" used in the description of the present application indicates that the elements or objects appearing before the word cover the elements or objects listed after the word and their equivalents, and other elements or objects are not excluded.

[0062] In addition, it should be further noted that, unless otherwise specified and limited, the "installation", "connection" and similar words used in the description of the present application should be understood in a broad sense, for example, the connection can be fixed connection, or detachable connection; can be mechanical connection, or electrical connection; can be directly connected, or indirectly connected through intermediate medium, and the person skilled in the art can understand the specific meaning of the present application according to the specific circumstances.

[0063] An aircraft sonic boom experiment model, as shown in Figure 1 , includes a front fuselage model 1, a middle fuselage model with wings 2, a fuselage rear body model with tail 3, a fuselage rear body model without tail 4, 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 with the front end of the middle fuselage model with wings 2. Specifically, the rear end of the front fuselage model 1 can be designed to have a connecting lug, the front end of the middle fuselage model with wings 2 can be designed to have a connecting lug hole, the connecting lug is inserted into the connecting lug hole, and a bolt is used for connection. The bolt can be arranged at 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, as shown in Figure 2 .

[0065] The tail winged fuselage rear body model 3 or the tail wingless fuselage rear body model 4 is detachably connected with the front end of the winged middle fuselage model 2, and the front end of the tail winged fuselage rear body model 3 or the tail wingless fuselage rear body model 4 is connected to the rear end of the winged middle fuselage model 2 according to the actual test setting. Specifically, the tail winged fuselage rear body model 3 or the tail wingless fuselage rear body model 4 is provided with a connecting column at the front end, and the winged middle fuselage model 2 is provided with a connecting hole at the rear end. When the tail winged fuselage rear body model 3 is connected with the winged middle fuselage model 2, the connecting column at the rear end of the tail winged fuselage rear body model 3 is inserted into the connecting hole, and the tail winged fuselage rear body model 3 is connected by a bolt. The bolt can be arranged at the bottom of the winged middle fuselage model 2, and the head of the bolt is flush with the outer surface of the winged middle fuselage model 2, as shown in FIG. 2. When the tail wingless fuselage rear body model 4 is connected with the winged middle fuselage model 2, the connecting column at the rear end of the tail wingless fuselage rear body model 4 is inserted into the connecting hole, and the tail wingless fuselage rear body model 4 is connected by a bolt. The bolt can be arranged at the bottom of the winged middle fuselage model 2, and the head of the bolt is flush with the outer surface of the winged middle fuselage model 2, as shown in FIG. 3. Figure 3 Figure 4

[0066] The front end of the support rod 5 is detachably connected to the back of the winged middle fuselage model 2, and the rear end of the support rod 5 is connected to the support structure. Specifically, the winged middle fuselage model 2 is provided with a strip-shaped hole at the back, and the support rod 5 is provided with a strip-shaped block at the front end. The strip-shaped block is arranged in the strip-shaped hole, and the support rod 5 is connected by a bolt. In addition, two blocking strips 6 are arranged in the strip-shaped hole, and the two blocking strips 6 are flush with the support rod 5, as shown in FIG. 4. The two blocking strips 6 and the heads of the bolts are connected to the winged middle fuselage model 2 by bolts, and the two blocking strips 6 and the heads of the bolts are flush with the outer surface of the winged middle fuselage model 2. Figure 5

[0067] The rear end of the front fuselage model 1 is provided with a mounting hole, and the blocking block 7 or the horizontal measurement platform 8 is arranged in the mounting hole. Specifically, when the aircraft sonic boom experimental model is installed and debugged, the horizontal measurement platform is arranged in the mounting hole to observe the angle of attack of the aircraft sonic boom experimental model, as shown in FIG. 5. When the sonic boom wind tunnel test is measured, the blocking block 7 is arranged in the mounting hole to fill the mounting hole, as shown in FIG. 6. Figure 6 Figure 7

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

[0069] Based on the aircraft sonic boom experimental model disclosed in the above embodiment, the application further provides a sonic boom wind tunnel test system, as shown in FIG. 7. Figure 8 ​​​​​As shown, including the aircraft sonic boom experiment model, model motion posture adjustment mechanism, sonic boom detection device, data acquisition and processing module.

[0070] Sonic boom wind tunnel test is an important means to carry out supersonic aircraft sonic boom research, limited by the size of the wind tunnel, can not reach the size of the direct simulation of far field, mainly for the measurement of near field out-of-body pressure distribution, the measured pressure spatial distribution, as the input condition of far field propagation model, and then get the far field sonic boom overpressure.

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

[0072] The model motion posture adjustment mechanism can be specifically designed to include guide rail 9, motion support 10, pitch adjustment block 11, suspension rod 12, as shown in Figure 9 .

[0073] The guide rail 9 is connected at the bottom of the wind tunnel, in order to reduce the influence on the wind tunnel, the guide rail 9 can be designed to have a small thickness.

[0074] The motion support 10 is connected to the guide rail 9 and can slide axially along the guide rail 9. Specifically, an actuator cylinder can be arranged between the motion support 10 and the guide rail 9, which can drive the motion support 10 to slide axially along the guide rail 9 by the extension and retraction of the piston rod, and further, the control end of the actuator cylinder can be connected to the wind tunnel measurement and control system, 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 size of the wind speed in the wind tunnel.

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

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

[0077] The moving support 10 slides along the guide rail 9 in the axial direction, and can drive the aircraft sonic boom experimental model to move close to or away from the sonic boom detection device in the axial direction. The pitch adjusting block 11 can be rotated to adjust the pitch attitude of the aircraft sonic boom experimental model, and change the angle of attack of the aircraft sonic boom experimental model.

[0078] The sonic boom detection device includes a plurality of sound pressure measuring points arranged in the axial direction 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 with a profile height of 350 mm and a top rounded portion. The top portion is arranged along the axis and has a total of 404 sound pressure measuring points with an inner diameter of 0.4 mm. The theoretical distance between adjacent sound pressure measuring points is 4 mm. The pressure measurement uses an Optimus system, and the module range includes 5 psid, 15 psid, 45 psid, and 75 psid. The full-scale accuracy is 0.05%.

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

[0080] The sonic boom wind tunnel test is essentially a measurement of the near-field spatial pressure distribution of the test model. Generally, 1-3 times the length of the model is selected. Such tests have high requirements for the flow field quality. In the supersonic wind tunnel, the model size is small due to the space limitation of the test section of the wind tunnel. The spatial non-uniformity and temporal non-stationarity of the wind tunnel flow field have a great influence on the measurement accuracy of the test. In addition, the support interference, model vibration, shock wave reflection, boundary layer interference, and other factors also affect the measurement accuracy of the wind tunnel test. Therefore, comprehensive consideration needs to be given in the test design, testing, and data processing.

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

[0082] The measurement vehicle data collection is carried out. The aircraft sonic boom experimental model is placed above the sound pressure measuring points, and the sonic boom wind tunnel test is carried out to measure the initial spatial pressure distribution data of the aircraft sonic boom experimental model and all components such as the track 9 in the flow field.

[0083] The reference vehicle data collection is carried out. The aircraft sonic boom experimental model is moved out of the measurement area or removed, and the reference spatial pressure distribution data of only the track 9 in the flow field is measured.

[0084] The initial spatial pressure distribution data is subtracted from the reference spatial pressure distribution data to obtain the spatial pressure change caused by the aircraft sonic boom experimental model, i.e. the near-field sonic boom overpressure distribution.

[0085] The reference train method assumes that the wind tunnel itself has no pressure fluctuation, and the net pressure distribution of the aircraft sonic boom experimental model is obtained by interference such as the rail 9. Due to the existence of pressure disturbance 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 disturbance of the wind tunnel flow field on the sonic boom signal measurement, a spatial averaging technique is further used to improve the measurement accuracy of the sonic boom test.

[0086] When measuring the train data, the aircraft sonic boom experimental model is controlled to move above the sound pressure measuring point along the axial direction at a certain interval distance, N positions are measured, and the initial spatial pressure distribution data and the reference spatial pressure distribution data are collected at each position. The spatial pressure change caused by the aircraft sonic boom experimental model is obtained by processing according to the following formula, and the spatial pressure change caused by the aircraft sonic boom experimental model is obtained by aligning the position and performing arithmetic average. The result is the spatially averaged near-field pressure distribution caused by the aircraft sonic boom experimental model.

[0087]

[0088] wherein,

[0089] is the spatial pressure change caused by the aircraft sonic boom experimental model;

[0090] is the overpressure of the aircraft sonic boom experimental model in the measurement area;

[0091] is the overpressure of the aircraft sonic boom experimental model outside the pressure measuring area;

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

[0093] The aircraft sonic boom experimental model and the sonic boom wind tunnel test system and method disclosed in the above embodiment are used for sonic boom test, and are convenient to assemble and adjust the shape. The test measurement efficiency and precision are high, and the test cost can be greatly reduced.

[0094] In one specific example, to verify the stability of the entire test system, a full mechanism type near-field sonic boom signal repeatability test is performed on the processed aircraft sonic boom experimental model. The test Mach number is 1.6, the model angle of attack is 3.5°, and the sonic boom signal measurement is at the position 1 times the body length below the model. The test result is as follows Figure 10The test measurement steps of train 1 and train 2 are 8, and the interval between steps is 4mm. The test data is the result after spatial averaging and zero drift correction. It can be seen that the overall law of the sound blast signal is basically the same, the sound blast signal law generated by the wing part of the two train tests is basically the same, the test data curves are well matched, and the maximum deviation is not more than 0.003. It shows that the test stability of the sound blast measurement system disclosed in the application is good, the test result deviation is small, and the need of the aircraft sound blast test can be met.

[0095] The test Mach number is 1.6, the model angle of attack is 3.5°, the sound blast signal measurement is at the position of 1 times the body length below the model, the full mechanism type sound blast test is carried out, and the sound blast signal is compared with the CFD calculation, as shown in Figure 11 The measurement steps of the wind tunnel test are 8, and the interval between steps is 4mm. The data is spatially averaged and corrected for zero drift. It can be seen that the overall law of the sound blast signal is basically the same, the sound blast signal law generated by the wing part of the two train tests is basically the same, the test data curves are well matched, and the maximum deviation is not more than 0.003. It shows that the test stability of the sound blast measurement system disclosed in the application is good, the test result deviation is small, and the need of the aircraft sound blast test can be met.

[0096] In order to further analyze the influence of support interference, the near-field sound blast signal of the aircraft sound blast experimental model with support is calculated by CFD, and compared with the wind tunnel test value. The flow field characteristics calculated by CFD are shown in Figure 12 The comparison of the sound blast test sound blast signal is shown in Figure 13 It can be seen from the calculation result that the numerical simulation result of the model with support is well matched with the wind tunnel test result, the peak value difference of the wave signal is 0.003 to 0.004, and the calculated value of the afterbody wave system has a certain difference with the test value. It shows that the support changes the pressure distribution near the wing, and then increases the peak value of the near-field overpressure distribution. It is necessary to strengthen the processing of support interference in the test result.

[0097] Based on an aircraft sonic boom test model, a sonic boom wind tunnel test was conducted on a tailless wing-body assembly. The test Mach number was 1.6, the model angle of attack was 3.5°, and the sonic boom signal was measured directly below the model at a position equal to one body length. The sonic boom signals of the full-body configuration and the tailless wing-body assembly configuration were compared. Figure 14 As shown, the measurement steps of the tailless wing-body combination wind tunnel test model are 5, with a 4mm interval between steps. The data are spatially averaged and zero-drift corrected. It can be seen that the sonic boom signal in the x / l = 0.9 to 1.77 range generally matches well, with only some differences in local locations. The changes in the sonic boom signal in the x / l = 1.77 to 1.9 range reflect the influence of the vertical tail and the high horizontal tail on the sonic boom signal. The test data shows that the tailless wing-body combination configuration has a larger absolute value of the negative peak value of the signal in this range, that is, stronger airflow expansion. This is consistent with the idea in the aerodynamic layout of supersonic aircraft with a rear-body sonic boom design, where the leading-edge shock wave generated by the high horizontal tail is more likely to merge with the wing trailing-edge expansion wave during propagation, so as to reduce the intensity of the sonic boom. In terms 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] Based on the sonic boom wind tunnel test of the tailless wing-body combination, sonic boom wind tunnel tests considering the influence of angle of attack were conducted by changing the incoming airflow angle. The test Mach number was 1.6, and the model angle of attack was 3.5° and 4.5°. The sonic boom signal was measured at a position one body length directly below the model. The sonic boom signal comparison curves for the tailless wing-body combination configuration are shown below. Figure 15 As shown, the wind tunnel test with a model angle of attack of 4.5° had 5 measurement steps with a 4mm interval between steps. The data were spatially averaged and corrected for zero drift. It can be seen that as the angle of attack increases, the sonic boom signal generally shows a "compression" trend, the sonic boom signal influence area shortens, 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 to 1.48 increases from 0.0169 to 0.0204, a change of about 0.003. In terms 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 angle of attack has a significant contribution to the lift coefficient.

[0099] The comprehensive test results show that the sonic boom wind tunnel test results can well reflect the distribution pattern of sonic boom signals of different models such as the whole aircraft and the 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] The technical scheme of the present application has been described in combination with the preferred embodiments shown in the drawings, and it should be understood by those skilled in the art that the protection scope of the present application is obviously not limited to these specific embodiments, and those skilled in the art can make equivalent changes or replacements to the related technical features without departing from the principles of the present application, and the technical schemes after the changes or replacements will fall within the protection scope of the present application.

Claims

1. A sonic boom wind tunnel testing 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 boom detection device is installed at the bottom of the wind tunnel; The aircraft sonic boom test model includes a forward fuselage model (1), a mid-fuselage model with wings (2), a rear fuselage model with tail fins (3), a rear fuselage model without tail fins (4), a support rod (5), a block (7), and a horizontal measurement platform (8). The rear end of the forward fuselage model (1) and the front end of the mid-fuselage model (2) with wings are detachably connected; The front end of the tail-equipped rear fuselage model (3) or the tailless rear fuselage model (4) is detachably connected to the rear end of the winged mid-fuselage model (2). The front end of the strut (5) is detachably connected to the back of the mid-fuselage model (2) with wings, and the rear end is connected to the support structure; The front fuselage model (1) has mounting holes at the rear end. A block (7) or a horizontal measuring platform (8) is installed in the mounting holes. When installing and debugging the aircraft sonic boom test model, the horizontal measuring platform is installed in the mounting holes to observe the angle of attack of the aircraft sonic boom test model. When conducting sonic boom wind tunnel test measurements, the block (7) is installed in the mounting holes to fill the mounting holes. The aircraft sonic boom experimental model is connected to the model motion adjustment mechanism through the rear end of the support rod (5). The aircraft sonic boom experimental model is located above the sonic boom detection device, at a distance of 1-3 times the model length from the top of the sonic boom detection device. The model motion adjustment mechanism can drive the aircraft sonic boom test model to move along the axial direction, move closer to or further 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 acquire and process near-field spatial pressure distribution measurement data of the aircraft sonic boom experimental model.

2. The sonic boom wind tunnel testing system according to claim 1, characterized in that, The front fuselage model (1) has connecting lugs at the rear end, and the mid-fuselage model (2) with wings has connecting lugs at the front end. The connecting lugs are inserted into the connecting lugs and connected by bolts. The bolts are located on the side of the mid-fuselage model (2) with wings, and their heads are flush with the outer surface of the mid-fuselage model (2) with wings.

3. The sonic boom wind tunnel testing system according to claim 2, characterized in that, The rear fuselage model with tail fin (3) and the rear fuselage model without tail fin (4) have connecting columns at the front end, and the mid-fuselage model with wings (2) has connecting holes at the rear end; When the tail-tailed rear fuselage model (3) is connected to the winged mid-fuselage model (2), the connecting post at the rear end of the tail-tailed rear fuselage model (3) is inserted into the connecting hole and connected by bolts. The bolts are set at the bottom of the winged mid-fuselage model (2) and the head is flush with the outer surface of the winged mid-fuselage model (2). When the tailless rear fuselage model (4) is connected to the winged mid-fuselage model (2), the connecting post at the rear end of the tailless rear fuselage model (4) is inserted into the connecting hole and connected by bolts. The bolts are located at the bottom of the winged mid-fuselage model (2) and the head is flush with the outer surface of the winged mid-fuselage model (2).

4. The sonic boom wind tunnel testing system according to claim 3, characterized in that, The mid-fuselage model with wings (2) has a strip hole on the back and a strip block at the front end of the support rod (5). The strip block is set in the strip hole and connected by bolts. Two plugs (6) are set in the strip hole. Two plugs (6) are attached to the support rod (5) and connected to the mid-fuselage model (2) with wings by bolts, and the two plugs (6) and their bolt heads are flush with the outer surface of the mid-fuselage model (2) with wings.

5. The sonic boom wind tunnel testing system according to claim 4, characterized in that, A force balance can be installed at the rear end of the support rod (5).

6. The sonic boom wind tunnel testing system 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 mounting surface is 5 degrees. The angle between the front mounting surface of the support rod (5) and the horizontal plane of the aircraft sonic boom experimental model is 2.5 degrees; The angle between the horizontal plane of the aircraft sonic boom experimental model and the rear axis of the support rod (5) is 2.5 degrees.

7. The sonic boom wind tunnel testing system according to claim 6, characterized in that, The model motion attitude adjustment mechanism includes a guide rail (9), a motion support (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 motion support (10) is connected to the guide rail (9) and can slide along the axial direction of the guide rail (9); The pitch adjustment block (11) is hinged to the motion support (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).

8. The sonic boom wind tunnel testing system according to claim 7, characterized in that, An actuator is installed between the motion support (10) and the guide rail (9). The actuator can drive the motion support (10) to slide along the guide rail (9) axially through the extension and retraction of the piston rod. The control end of the actuator is connected to the wind tunnel measurement and control system. The extension and retraction of its 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 inside the wind tunnel.

9. The sonic boom wind tunnel testing system according to claim 8, characterized in that, An actuator is provided between the pitch adjustment block (11) and the motion support (10). The actuator can drive the pitch adjustment block (11) to rotate by the extension and retraction of the piston rod. The control end of the actuator is connected to the wind tunnel measurement and control system, and the extension and retraction of its piston rod is controlled by the wind tunnel measurement and control system.

10. The sonic boom wind tunnel testing system according to claim 9, characterized in that, The sonic boom detection device includes multiple sonic pressure measuring points set along the axial direction at the front end of the guide rail (9).

11. The sonic boom wind tunnel testing system according to claim 10, characterized in that, The guide rail (9) has a measuring section (13) at the front end. The measuring section has a cross-sectional height of 350mm and a rounded top. A total of 404 sound pressure measuring points with an inner diameter of 0.4mm are arranged along the axis at the top. The theoretical distance between adjacent sound pressure measuring points is 4mm.

12. A sonic boom wind tunnel testing method, implemented based on the sonic boom wind tunnel testing system described in claim 11, characterized in that, include: Data collection was carried out for the measurement trains. The aircraft sonic boom experimental model was placed above the sound pressure measurement point, and a sonic boom wind tunnel test was conducted to measure the initial space pressure distribution data. Reference train data collection was carried out, and the aircraft sonic boom experimental model was moved outside the measurement area or dismantled to obtain reference space pressure distribution data. By subtracting the reference spatial pressure distribution data from the initial spatial pressure distribution data, the spatial pressure change caused by the aircraft sonic boom experimental model is obtained, i.e., the near-field sonic boom overpressure distribution.

13. The sonic boom wind tunnel testing method according to claim 12, characterized in that, When collecting data for measurement trains, the aircraft sonic boom experimental model is controlled to move along the axial direction at certain intervals above the sound pressure measurement point, and N positions are measured. At each position, initial spatial pressure distribution data and reference spatial pressure distribution data are collected once to obtain N sets of spatial pressure changes caused by the aircraft sonic boom experimental model. After aligning the positions of these N sets of spatial pressure changes caused by the aircraft sonic boom experimental model, an arithmetic average is taken, and the result is the spatially averaged spatial pressure change caused by the aircraft sonic boom experimental model.

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