Pub Date : 2024-06-14DOI: 10.1177/1475472x241259082
Seçkin Uluskan
This study introduces a new method for precise acoustic drone localization and tracking via the noise generated by the drone. Drone noises include harmonic frequency components which are related to the rotational speed of the propeller and the number of blades. This study integrates utilization of the frequency components around harmonics into Steered Response Power - Phase Transform (SRP-PHAT). First, a custom discrete Fourier transform (namely DFT-Harmonics) is defined which concentrates only on the vicinities of harmonics to capture the frequency components possibly related to the drone sound. Then, DFT-Harmonics is integrated into SRP-PHAT, which is named SRP-Harmonics. The benefits of SRP-Harmonics are explained and illustrated via SRP maps and videos. Experiments with real microphone array data show that SRP-Harmonics is precise in localizing and tracking a drone, while the ordinary SRP-PHAT can not be reasonably successful. Moreover, SRP-Harmonics after Kaiser window can maintain its performance even when significant level of artificial white noise or natural wind noise exists in the data.
{"title":"Precise acoustic drone localization and tracking via drone noise: Steered response power - phase transform around harmonics","authors":"Seçkin Uluskan","doi":"10.1177/1475472x241259082","DOIUrl":"https://doi.org/10.1177/1475472x241259082","url":null,"abstract":"This study introduces a new method for precise acoustic drone localization and tracking via the noise generated by the drone. Drone noises include harmonic frequency components which are related to the rotational speed of the propeller and the number of blades. This study integrates utilization of the frequency components around harmonics into Steered Response Power - Phase Transform (SRP-PHAT). First, a custom discrete Fourier transform (namely DFT-Harmonics) is defined which concentrates only on the vicinities of harmonics to capture the frequency components possibly related to the drone sound. Then, DFT-Harmonics is integrated into SRP-PHAT, which is named SRP-Harmonics. The benefits of SRP-Harmonics are explained and illustrated via SRP maps and videos. Experiments with real microphone array data show that SRP-Harmonics is precise in localizing and tracking a drone, while the ordinary SRP-PHAT can not be reasonably successful. Moreover, SRP-Harmonics after Kaiser window can maintain its performance even when significant level of artificial white noise or natural wind noise exists in the data.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-06-14","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141338571","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-06-05DOI: 10.1177/1475472x241259101
Hongjun Qian, Yusen Niu, Yi Jiang, Peize Yan
Rocket sled test avoids boundary effects in wind tunnel test and inconvenience of flight test, which is one of the feasible options for future single-stage-to-orbit. To analyze potential safety issues during payload separation and optimize the arrangement of testing sensors, different structural layouts and operating speeds of the rocket sled are conducted based on computational fluid dynamics and computational aeroacoustics. The hypersonic winged standard model is utilized as the load for these simulations. The analysis encompasses the evolution of shock waves, the forces exerted on the payload and noise propagation. Variations in flow field and aeroacoustic characteristics of rocket sled are analysed, revealing underlying physical mechanisms. It is observed that placing the payload in front of the thruster can reduce the head pressure, excessive wingspan may have an impact on the structural safety of the wingtip, and the regions of high sound pressure levels mainly exists in the middle and rear sections of the rocket sled. Moreover, as the Mach number increases, the characteristic frequency initially rises and then declines. These researches can serve as a valuable reference for the development of ground test systems and single-stage-to-orbit.
{"title":"Aerodynamic and aeroacoustic characteristics of rocket sled under strong ground effect","authors":"Hongjun Qian, Yusen Niu, Yi Jiang, Peize Yan","doi":"10.1177/1475472x241259101","DOIUrl":"https://doi.org/10.1177/1475472x241259101","url":null,"abstract":"Rocket sled test avoids boundary effects in wind tunnel test and inconvenience of flight test, which is one of the feasible options for future single-stage-to-orbit. To analyze potential safety issues during payload separation and optimize the arrangement of testing sensors, different structural layouts and operating speeds of the rocket sled are conducted based on computational fluid dynamics and computational aeroacoustics. The hypersonic winged standard model is utilized as the load for these simulations. The analysis encompasses the evolution of shock waves, the forces exerted on the payload and noise propagation. Variations in flow field and aeroacoustic characteristics of rocket sled are analysed, revealing underlying physical mechanisms. It is observed that placing the payload in front of the thruster can reduce the head pressure, excessive wingspan may have an impact on the structural safety of the wingtip, and the regions of high sound pressure levels mainly exists in the middle and rear sections of the rocket sled. Moreover, as the Mach number increases, the characteristic frequency initially rises and then declines. These researches can serve as a valuable reference for the development of ground test systems and single-stage-to-orbit.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-06-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141383827","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-06-04DOI: 10.1177/1475472x241259103
R. Matouk
The accurate prediction of the trailing-edge noise and the determination of its sources are essential to reduce fans and propellers noise. This noise component is due to the scattering of the turbulent boundary layer into acoustic waves by the trailing edge. In this paper, the noise emanating from a CD (Controlled-Diffusion) airfoil is simulated and computed via the hybrid methods of aeroacoustics. In these methods, the aerodynamic and acoustic fields are computed separately. The flow data are obtained using the in-house LES solver SFELES. ACTRAN acoustic solver has been used to solve the acoustics and to provide the near and far fields propagation via Lighthill’s analogy. Curle’s analogy is applied as well in its integral compact formulation which takes the presence of walls into account. Curle’s formulation is applied proposing an approach where the volume and surface integrals have been implemented in SFELES to be calculated simultaneously with the flow in order to avoid the storage of noise sources which requires a huge space. In Lighthill’s analogy, sources and near field maps show that the turbulent boundary layer and wake are the more efficient sources and the center of radiation is the trailing edge. The comparison of the numerical results with the experimental measurements, performed by Moreau and Roger and Moreau et al. , shows an overall excellent agreement confirming the capability of SFELES (LES sources) combined with ACTRAN (Lighthill’s analogy) to predict correctly the noise generated by turbulent flows around airfoils. The acoustic spectrum presents an overprediction up to 5 dB at the frequencies 300 Hz and 550 Hz and an underprediction about 5 dB at the frequencies 1100 Hz and 1750 Hz. The sound pressure level (SPL) obtained using the proposed approach of Curle’s analogy matches very well the experimental results. Thus, Curle’s analogy can be used to obtain a fast, approximated and acceptable results about the noise radiation of airfoils avoiding the storage of noise sources which requires a huge space and time.
{"title":"Prediction of the aerodynamic noise of an airfoil via the hybrid methods of aeroacoustics","authors":"R. Matouk","doi":"10.1177/1475472x241259103","DOIUrl":"https://doi.org/10.1177/1475472x241259103","url":null,"abstract":"The accurate prediction of the trailing-edge noise and the determination of its sources are essential to reduce fans and propellers noise. This noise component is due to the scattering of the turbulent boundary layer into acoustic waves by the trailing edge. In this paper, the noise emanating from a CD (Controlled-Diffusion) airfoil is simulated and computed via the hybrid methods of aeroacoustics. In these methods, the aerodynamic and acoustic fields are computed separately. The flow data are obtained using the in-house LES solver SFELES. ACTRAN acoustic solver has been used to solve the acoustics and to provide the near and far fields propagation via Lighthill’s analogy. Curle’s analogy is applied as well in its integral compact formulation which takes the presence of walls into account. Curle’s formulation is applied proposing an approach where the volume and surface integrals have been implemented in SFELES to be calculated simultaneously with the flow in order to avoid the storage of noise sources which requires a huge space. In Lighthill’s analogy, sources and near field maps show that the turbulent boundary layer and wake are the more efficient sources and the center of radiation is the trailing edge. The comparison of the numerical results with the experimental measurements, performed by Moreau and Roger and Moreau et al. , shows an overall excellent agreement confirming the capability of SFELES (LES sources) combined with ACTRAN (Lighthill’s analogy) to predict correctly the noise generated by turbulent flows around airfoils. The acoustic spectrum presents an overprediction up to 5 dB at the frequencies 300 Hz and 550 Hz and an underprediction about 5 dB at the frequencies 1100 Hz and 1750 Hz. The sound pressure level (SPL) obtained using the proposed approach of Curle’s analogy matches very well the experimental results. Thus, Curle’s analogy can be used to obtain a fast, approximated and acceptable results about the noise radiation of airfoils avoiding the storage of noise sources which requires a huge space and time.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-06-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141387603","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-06-02DOI: 10.1177/1475472x241259122
Weicheng Xue, Bing Yang
The deconvolution DAMAS algorithm can effectively eliminate the misconceptions in the usually-used beamforming localization algorithm, allowing for a more accurate calculation of the source location as well as the intensity. When solving a linear system of equations, the DAMAS algorithm takes into account the mutual influence of different locations, reducing or even eliminating sidelobes and producing more accurate results. This work first introduces the principles of the DAMAS algorithm. Then it applies both the conventional beamforming algorithm and the DAMAS algorithm to simulate the localization of a single-frequency source from a 1.5 MW wind turbine, a complex line source with the text “UCAS” and a line source downstream of an airfoil trailing edge. Finally, the work presents experimental localization results of the source of a 1.5 MW wind turbine using both the conventional beamforming algorithm and the DAMAS algorithm.
{"title":"Aeroacoustic source localization using the microphone array method with application to wind turbine noise","authors":"Weicheng Xue, Bing Yang","doi":"10.1177/1475472x241259122","DOIUrl":"https://doi.org/10.1177/1475472x241259122","url":null,"abstract":"The deconvolution DAMAS algorithm can effectively eliminate the misconceptions in the usually-used beamforming localization algorithm, allowing for a more accurate calculation of the source location as well as the intensity. When solving a linear system of equations, the DAMAS algorithm takes into account the mutual influence of different locations, reducing or even eliminating sidelobes and producing more accurate results. This work first introduces the principles of the DAMAS algorithm. Then it applies both the conventional beamforming algorithm and the DAMAS algorithm to simulate the localization of a single-frequency source from a 1.5 MW wind turbine, a complex line source with the text “UCAS” and a line source downstream of an airfoil trailing edge. Finally, the work presents experimental localization results of the source of a 1.5 MW wind turbine using both the conventional beamforming algorithm and the DAMAS algorithm.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-06-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"141273213","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230648
Nikolai Pastouchenko, Kishore Ramakrishnan, Steven Morris, Umesh Paliath
Jet noise generated by turbofan engines which are designed for supersonic aircraft operation is expected to remain a challenge for the aerospace industry. For civil supersonic transport, low bypass ratio engines with internal mixing and external plugs are currently seen as promising exhaust architecture to meet aircraft mission and certification requirements. Measurements of the noise from a representative low-bypass ratio long duct mixed flow with external plug exhaust configuration were made, along with corresponding steady-RANS and LES analyses. The measured data showed excess high frequency noise across a wide range of operating conditions as compared to an equivalent single stream jet, or an equivalent uniform jet with internal plug. The absolute levels of the excess noise were relatively insensitive to forward flight, and the excess noise was most dominant when the primary and secondary streams had large velocity and temperature mismatch just upstream of the internal mixer. Preliminary indications from the LES analysis also indicated that significant pressure fluctuations, at the same frequencies which are insensitive to forward flight, develop inside the nozzle as the internal shear layer from the mixer grows and gets accelerated by the external plug. Parametric study test measurements on the fan and core operating conditions were performed, providing support for the hypothesis that the excess noise is generated as the internal shear layer, starting inside the nozzle, gets accelerated by the plug and passes through the weak shock at the nozzle exit where the flow accelerates more and turns over the external plug.
为超音速飞机运行而设计的涡轮风扇发动机产生的喷气噪声预计仍将是航空航天工业面临的一项挑战。对于民用超音速运输来说，具有内部混合和外部塞子的低旁通比发动机目前被视为有希望满足飞机任务和认证要求的排气结构。我们测量了具有代表性的低旁通比长管道混流外塞排气结构的噪声，并进行了相应的稳定-RANS 和 LES 分析。测量数据显示，与等效的单流喷气机或等效的带内塞的均匀喷气机相比，高频噪声在很大的工作条件范围内都超标。过量噪声的绝对值对正向飞行相对不敏感，当主流和副流在内部混合器上游有较大的速度和温度不匹配时，过量噪声最为显著。LES 分析的初步结果还表明，随着来自混合器的内部剪切层的增长和外部塞子的加速，喷嘴内部也会产生显著的压力波动，其频率与对前向飞行不敏感的频率相同。对风扇和核心筒的工作条件进行了参数研究测试测量，为以下假设提供了支持：当内部剪切层从喷嘴内部开始，被塞子加速并通过喷嘴出口处的弱冲击时，就会产生过大的噪声，在那里，气流加速度更大，并翻过外部塞子。
{"title":"Jet noise of a low bypass turbofan with internal mixing and external plug","authors":"Nikolai Pastouchenko, Kishore Ramakrishnan, Steven Morris, Umesh Paliath","doi":"10.1177/1475472x241230648","DOIUrl":"https://doi.org/10.1177/1475472x241230648","url":null,"abstract":"Jet noise generated by turbofan engines which are designed for supersonic aircraft operation is expected to remain a challenge for the aerospace industry. For civil supersonic transport, low bypass ratio engines with internal mixing and external plugs are currently seen as promising exhaust architecture to meet aircraft mission and certification requirements. Measurements of the noise from a representative low-bypass ratio long duct mixed flow with external plug exhaust configuration were made, along with corresponding steady-RANS and LES analyses. The measured data showed excess high frequency noise across a wide range of operating conditions as compared to an equivalent single stream jet, or an equivalent uniform jet with internal plug. The absolute levels of the excess noise were relatively insensitive to forward flight, and the excess noise was most dominant when the primary and secondary streams had large velocity and temperature mismatch just upstream of the internal mixer. Preliminary indications from the LES analysis also indicated that significant pressure fluctuations, at the same frequencies which are insensitive to forward flight, develop inside the nozzle as the internal shear layer from the mixer grows and gets accelerated by the external plug. Parametric study test measurements on the fan and core operating conditions were performed, providing support for the hypothesis that the excess noise is generated as the internal shear layer, starting inside the nozzle, gets accelerated by the plug and passes through the weak shock at the nozzle exit where the flow accelerates more and turns over the external plug.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140996825","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230647
Mathieu Varé, C. Bogey
The differences between the acoustic tones generated by impinging jets with laminar and highly-disturbed nozzle-exit boundary layers are investigated. For that, jets at Mach numbers between 0.6 and 1.3 impinging on a flat plate at a distance of 8 nozzle radii from the nozzle exit are computed using large-eddy simulations. The amplitudes of the tones generated by the jets through feedback loops establishing between the nozzle and the plate are found to be significantly affected by the exit turbulent disturbances. In the present study, overall, they are lower for the initially laminar jets than for the initially disturbed ones. The level decrease varies from a few dB up to 15 dB, depending on the tones, which can change the frequencies of the dominant tones and the numbers and azimuthal structures of their associated feedback modes. For Mach numbers 0.75 and 0.8, for instance, the dominant tone frequencies are approximately two times lower for the initially laminar jets than for the other ones, yielding a better agreement with experiments of the literature in the former case. For a Mach number of 1.1, as a second example, the dominant tone is associated with the axisymmetric third feedback mode in the laminar case but with the helical fifth feedback mode in the disturbed case. The differences in the tone amplitude are finally discussed by estimating the power gains of the shear-layer instability waves between the nozzle and the plate using linear stability analysis for the axisymmetric mode. In most cases, at the frequency of a specific tone, the higher the gain, the stronger the acoustic tone.
研究了具有层流边界层和高扰动喷嘴出口边界层的撞击喷流所产生的声调之间的差异。为此，使用大涡流模拟计算了马赫数在 0.6 和 1.3 之间的喷流撞击到距离喷嘴出口 8 个喷嘴半径处的平板上的情况。研究发现，喷流通过在喷嘴和平板之间建立的反馈回路产生的音调振幅受到出口湍流扰动的显著影响。在本研究中，总体而言，初始层流喷流的音调低于初始扰动喷流。电平降低幅度从几 dB 到 15 dB 不等，取决于音调，这可能会改变主要音调的频率及其相关反馈模式的数量和方位角结构。例如，在马赫数为 0.75 和 0.8 时，初始层流喷流的主音频率比其他喷流的主音频率低约两倍，前者与文献中的实验结果更为一致。第二个例子是马赫数为 1.1 时，层流情况下的主音与轴对称第三反馈模式有关，而扰动情况下的主音与螺旋第五反馈模式有关。通过对轴对称模式进行线性稳定性分析，估算喷嘴和板之间剪切层不稳定波的功率增益，最终讨论了音调振幅的差异。在大多数情况下，在特定声调的频率下，增益越高，声调越强。
{"title":"Acoustic tones generated by impinging jets: Differences between laminar and highly-disturbed nozzle-exit boundary layers","authors":"Mathieu Varé, C. Bogey","doi":"10.1177/1475472x241230647","DOIUrl":"https://doi.org/10.1177/1475472x241230647","url":null,"abstract":"The differences between the acoustic tones generated by impinging jets with laminar and highly-disturbed nozzle-exit boundary layers are investigated. For that, jets at Mach numbers between 0.6 and 1.3 impinging on a flat plate at a distance of 8 nozzle radii from the nozzle exit are computed using large-eddy simulations. The amplitudes of the tones generated by the jets through feedback loops establishing between the nozzle and the plate are found to be significantly affected by the exit turbulent disturbances. In the present study, overall, they are lower for the initially laminar jets than for the initially disturbed ones. The level decrease varies from a few dB up to 15 dB, depending on the tones, which can change the frequencies of the dominant tones and the numbers and azimuthal structures of their associated feedback modes. For Mach numbers 0.75 and 0.8, for instance, the dominant tone frequencies are approximately two times lower for the initially laminar jets than for the other ones, yielding a better agreement with experiments of the literature in the former case. For a Mach number of 1.1, as a second example, the dominant tone is associated with the axisymmetric third feedback mode in the laminar case but with the helical fifth feedback mode in the disturbed case. The differences in the tone amplitude are finally discussed by estimating the power gains of the shear-layer instability waves between the nozzle and the plate using linear stability analysis for the axisymmetric mode. In most cases, at the frequency of a specific tone, the higher the gain, the stronger the acoustic tone.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140997115","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230650
Richard Auhl, Dennis K. McLaughlin, Philip J. Morris
The Aeroacoustics Laboratory in the Department of Aerospace Engineering at Penn State currently serves several areas of experimental research. Most notable in the laboratory is the Anechoic Chamber of which the Aeroacoustics Wind Tunnel is an integral part. In recent years the dominant areas of research have included model aircraft exhaust jets and model rotors for helicopter applications involving detailed noise experiments. In the jet noise area, the applications have included aircraft in both flyover and take-off configurations. In the flyover application the air flow surrounding the exhaust jet has a significant effect on the radiated noise and must be simulated to produce experimental results that will be most useful in preliminary design of noise suppression applications. The Aeroacoustics Wind Tunnel at Penn State serves the purpose for this simulation. During aircraft take-off, the aircraft airspeed is in a range exceeding 250 ft per second, or Mach number of over M = 0.22. In the years preceding 2022, the velocity of the free jet flow of the Penn State facility was limited to less than 200 ft/sec. This paper reports on a project to design upgrades to the existing wind tunnel to improve the test section flow velocity to values closer to the 250 ft/sec target. The approach began with a preliminary design of the return flow ducting to convert the open jet, open return wind tunnel to a closed return tunnel. The concept is to make use of the flow energy in the exhaust of the tunnel to boost the input energy to the inlet fan that drives the flow to the test section. An integral part of the preliminary design involved making use of an upgraded flow analysis computer code to predict the gain in test section velocity of the new facility. For the available horsepower of two fans in the facility, the configurations of the existing and upgraded wind tunnel were entered into this analysis code based on a quasi-one-dimensional formulation. The code included empirical data formulated to include the various shapes of the components of the tunnels. The relative flow velocities in each section used simple incompressible continuity to calculate all velocities in the component sections. The average dynamic pressure in each section is calculated from the flow (constant) density and the velocity squared. Each section had a non-dimensional pressure loss parameter calculated from empirical data (from the relevant literature). The basic code (using the Excel algorithms) was based on a predictor-corrector concept in which the calculation was initialized with the estimated velocity and pressure of the flow at the entrance to the test section. The loss in each following section is estimated to determine the loss in the total pressure and the cumulated loss compared to the gains made at the two fans in the circuit. The initial guess is adjusted based on the total pressure value at the end of the circuit. During the related experiments, measurements of the total and
{"title":"Design, assembly and testing of an upgraded aeroacoustics wind tunnel facility","authors":"Richard Auhl, Dennis K. McLaughlin, Philip J. Morris","doi":"10.1177/1475472x241230650","DOIUrl":"https://doi.org/10.1177/1475472x241230650","url":null,"abstract":"The Aeroacoustics Laboratory in the Department of Aerospace Engineering at Penn State currently serves several areas of experimental research. Most notable in the laboratory is the Anechoic Chamber of which the Aeroacoustics Wind Tunnel is an integral part. In recent years the dominant areas of research have included model aircraft exhaust jets and model rotors for helicopter applications involving detailed noise experiments. In the jet noise area, the applications have included aircraft in both flyover and take-off configurations. In the flyover application the air flow surrounding the exhaust jet has a significant effect on the radiated noise and must be simulated to produce experimental results that will be most useful in preliminary design of noise suppression applications. The Aeroacoustics Wind Tunnel at Penn State serves the purpose for this simulation. During aircraft take-off, the aircraft airspeed is in a range exceeding 250 ft per second, or Mach number of over M = 0.22. In the years preceding 2022, the velocity of the free jet flow of the Penn State facility was limited to less than 200 ft/sec. This paper reports on a project to design upgrades to the existing wind tunnel to improve the test section flow velocity to values closer to the 250 ft/sec target. The approach began with a preliminary design of the return flow ducting to convert the open jet, open return wind tunnel to a closed return tunnel. The concept is to make use of the flow energy in the exhaust of the tunnel to boost the input energy to the inlet fan that drives the flow to the test section. An integral part of the preliminary design involved making use of an upgraded flow analysis computer code to predict the gain in test section velocity of the new facility. For the available horsepower of two fans in the facility, the configurations of the existing and upgraded wind tunnel were entered into this analysis code based on a quasi-one-dimensional formulation. The code included empirical data formulated to include the various shapes of the components of the tunnels. The relative flow velocities in each section used simple incompressible continuity to calculate all velocities in the component sections. The average dynamic pressure in each section is calculated from the flow (constant) density and the velocity squared. Each section had a non-dimensional pressure loss parameter calculated from empirical data (from the relevant literature). The basic code (using the Excel algorithms) was based on a predictor-corrector concept in which the calculation was initialized with the estimated velocity and pressure of the flow at the entrance to the test section. The loss in each following section is estimated to determine the loss in the total pressure and the cumulated loss compared to the gains made at the two fans in the circuit. The initial guess is adjusted based on the total pressure value at the end of the circuit. During the related experiments, measurements of the total and ","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140995835","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230654
K. Zaman, D. J. Dicki
A scale model experimental study is conducted addressing howl occurring during engine tests in the Propulsion Systems Laboratory (PSL) of NASA Glenn Research Center. The aim is to understand the sources of the howling noise and find possible remedies. It is shown that a jet discharged into the model of the PSL entrance duct gives rise to high levels of low frequency noise. This occurs apparently due to duct acoustic modes randomly excited by the jet. Wire-mesh screens of various mesh size and configuration as well as tabs, placed at the end of the model duct, are shown to affect the noise. Comparative effects of these configurations are studied, and some are found to alleviate the low-frequency noise significantly. Based on these results, the equivalent of a 4-mesh screen is recommended as the end treatment for the duct of the PSL facility.
{"title":"A scale model study of an engine test facility for alleviating low-frequency howl","authors":"K. Zaman, D. J. Dicki","doi":"10.1177/1475472x241230654","DOIUrl":"https://doi.org/10.1177/1475472x241230654","url":null,"abstract":"A scale model experimental study is conducted addressing howl occurring during engine tests in the Propulsion Systems Laboratory (PSL) of NASA Glenn Research Center. The aim is to understand the sources of the howling noise and find possible remedies. It is shown that a jet discharged into the model of the PSL entrance duct gives rise to high levels of low frequency noise. This occurs apparently due to duct acoustic modes randomly excited by the jet. Wire-mesh screens of various mesh size and configuration as well as tabs, placed at the end of the model duct, are shown to affect the noise. Comparative effects of these configurations are studied, and some are found to alleviate the low-frequency noise significantly. Based on these results, the equivalent of a 4-mesh screen is recommended as the end treatment for the duct of the PSL facility.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140997586","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230644
K. Viswanathan, Fang Q Hu
This introduction provides a brief personal background and highlights the scientific accomplishments of Professor Christopher Tam. He has maintained extremely high standards of research and an enviable track record for productivity, spanning more than five decades. The depth and breadth of his technical output are truly breathtaking, and he has inspired scores of researchers and graduate students. When the first author was an Acoustics Engineer at Lockheed, Georgia, he started collaborating with Chris in the early 1990s. There were several joint projects when the author moved to Boeing, which encompassed both fundamental research and applications. The second author has an even longer association with Chris, initially as a PhD advisee in the late 1980s and then as a co-researcher. In this special volume we pay tribute to Chris’s many achievements, an endeavor that provides us with great satisfaction and which is long overdue.
{"title":"Christopher Tam: Brief history and accomplishments","authors":"K. Viswanathan, Fang Q Hu","doi":"10.1177/1475472x241230644","DOIUrl":"https://doi.org/10.1177/1475472x241230644","url":null,"abstract":"This introduction provides a brief personal background and highlights the scientific accomplishments of Professor Christopher Tam. He has maintained extremely high standards of research and an enviable track record for productivity, spanning more than five decades. The depth and breadth of his technical output are truly breathtaking, and he has inspired scores of researchers and graduate students. When the first author was an Acoustics Engineer at Lockheed, Georgia, he started collaborating with Chris in the early 1990s. There were several joint projects when the author moved to Boeing, which encompassed both fundamental research and applications. The second author has an even longer association with Chris, initially as a PhD advisee in the late 1980s and then as a co-researcher. In this special volume we pay tribute to Chris’s many achievements, an endeavor that provides us with great satisfaction and which is long overdue.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140995018","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-05-09DOI: 10.1177/1475472x241230649
Fang Q Hu, D. Nark
It has been well-known that the Ingard-Myers impedance condition, while simple to apply, is subject to the hydrodynamic Kelvin-Helmholtz-type instability due to its use of a vortex sheet in modeling the flow at the liner boundary. Recently, in the development of a time domain boundary element method for acoustic scattering by treated surfaces, it was found that by neglecting a certain second-order spatial derivative term in the Ingard-Myers formulation, the hydrodynamic instability can be avoided. The present paper aims to provide further analysis of this modified condition, hereby referred to as the Truncated Ingard-Myers Impedance Boundary Condition (TIMIBC). It will be shown, based on the dispersion relations of linear waves, that the instability intrinsic to the Ingard-Myers condition is eliminated in the proposed new formulation. Quantitative assessments on the accuracy of the TIMIBC for scattering of acoustic waves by lined surfaces are carried out, and its effectiveness is demonstrated by a numerical example. It is found that the TIMIBC provides a good approximation to the original Ingard-Myers condition for flows of low to mid subsonic Mach numbers. As such, the proposed TIMIBC can offer a practical solution for overcoming the intrinsic instability associated with the Ingard-Myers condition. Moreover, time domain implementation of the TIMIBC is also discussed and illustrated with a numerical example using a finite difference scheme. In particular, a minimization procedure for finding the poles and coefficients of a broadband multipole expansion for the impedance function is formulated by which, unlike the commonly used vector-fitting method, passivity of the model is ensured.
{"title":"On a stabilization of the Ingard-Myers impedance boundary condition and its time domain implementation","authors":"Fang Q Hu, D. Nark","doi":"10.1177/1475472x241230649","DOIUrl":"https://doi.org/10.1177/1475472x241230649","url":null,"abstract":"It has been well-known that the Ingard-Myers impedance condition, while simple to apply, is subject to the hydrodynamic Kelvin-Helmholtz-type instability due to its use of a vortex sheet in modeling the flow at the liner boundary. Recently, in the development of a time domain boundary element method for acoustic scattering by treated surfaces, it was found that by neglecting a certain second-order spatial derivative term in the Ingard-Myers formulation, the hydrodynamic instability can be avoided. The present paper aims to provide further analysis of this modified condition, hereby referred to as the Truncated Ingard-Myers Impedance Boundary Condition (TIMIBC). It will be shown, based on the dispersion relations of linear waves, that the instability intrinsic to the Ingard-Myers condition is eliminated in the proposed new formulation. Quantitative assessments on the accuracy of the TIMIBC for scattering of acoustic waves by lined surfaces are carried out, and its effectiveness is demonstrated by a numerical example. It is found that the TIMIBC provides a good approximation to the original Ingard-Myers condition for flows of low to mid subsonic Mach numbers. As such, the proposed TIMIBC can offer a practical solution for overcoming the intrinsic instability associated with the Ingard-Myers condition. Moreover, time domain implementation of the TIMIBC is also discussed and illustrated with a numerical example using a finite difference scheme. In particular, a minimization procedure for finding the poles and coefficients of a broadband multipole expansion for the impedance function is formulated by which, unlike the commonly used vector-fitting method, passivity of the model is ensured.","PeriodicalId":49304,"journal":{"name":"International Journal of Aeroacoustics","volume":null,"pages":null},"PeriodicalIF":1.0,"publicationDate":"2024-05-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"140995595","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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