In order to reduce oil dynamic power loss in aeroengine gearboxes, visualizations and measurements of the oil-flow are effective. In the research presented in this paper, we developed a flow visualization borescope which can qualitatively visualize oil flow and a two-phase flow probe which can quantitatively measure oil/air ratio and the flow velocity. The flow visualization borescope consists of a 16mm diameter pipe in which an air purge passage for removing oil mist and a borescope are integrated with an illumination laser light and optical lenses, enabling clear high-speed photography. The two-phase probe consists of a 5mm diameter pipe with a 1mm diameter measurement hole and has a pressure adjustment pipe inside the pipe. For a demonstration, a shrouded spur gear with 100 m/s peripheral speed and 20 liters/min oil supply was used. Flow visualization at 30000 frame/sec imaging shows that oil outflow from the shroud opening spreads turbulently over the whole width of the opening. Oil/air ratio and flow velocity measurement by the two-phase flow probe show that there was thin oil-rich layer on the shroud wall and the flow speed was slow compared with the gear peripheral speed. The measurement equipment we developed was easily installed to the gearbox and therefore it is expected to apply to real aeroengine gearboxes.
为了减少航空发动机齿轮箱中油流的动功率损失，油流的可视化和测量是有效的。在本文的研究中，我们开发了一种可以定性地显示油流的流动显示内窥镜和一种可以定量测量油气比和流速的两相流探头。流动可视化内窥镜由直径16mm的管道组成，管道中有一个用于去除油雾的空气吹扫通道，内窥镜与照明激光和光学透镜集成在一起，可以实现清晰的高速摄影。两相探头由直径5mm的管道和直径1mm的测量孔组成，管内有压力调节管。为进行验证，采用外围转速为100 m/s、供油量为20 l /min的带冠直齿齿轮。30000帧/秒成像下的流动显示，从叶冠开口流出的油在整个开口宽度上呈湍流扩散。两相流探头测量的油气比和流速表明，叶冠壁面上存在较薄的富油层，且流速较齿轮外周速度慢。我们开发的测量设备易于安装在变速箱上，因此有望应用于实际的航空发动机变速箱。
{"title":"Developments of a Flow Visualization Borescope and a Two-Phase Flow Probe for Aeroengine Transmission Gears","authors":"Hidenori Arisawa, Yuji Shinoda, Noguchi Yoshiyuki, T. Goi, Takahiko Banno, Hirofumi Akahori","doi":"10.1115/GT2018-75083","DOIUrl":"https://doi.org/10.1115/GT2018-75083","url":null,"abstract":"In order to reduce oil dynamic power loss in aeroengine gearboxes, visualizations and measurements of the oil-flow are effective. In the research presented in this paper, we developed a flow visualization borescope which can qualitatively visualize oil flow and a two-phase flow probe which can quantitatively measure oil/air ratio and the flow velocity. The flow visualization borescope consists of a 16mm diameter pipe in which an air purge passage for removing oil mist and a borescope are integrated with an illumination laser light and optical lenses, enabling clear high-speed photography. The two-phase probe consists of a 5mm diameter pipe with a 1mm diameter measurement hole and has a pressure adjustment pipe inside the pipe. For a demonstration, a shrouded spur gear with 100 m/s peripheral speed and 20 liters/min oil supply was used. Flow visualization at 30000 frame/sec imaging shows that oil outflow from the shroud opening spreads turbulently over the whole width of the opening. Oil/air ratio and flow velocity measurement by the two-phase flow probe show that there was thin oil-rich layer on the shroud wall and the flow speed was slow compared with the gear peripheral speed. The measurement equipment we developed was easily installed to the gearbox and therefore it is expected to apply to real aeroengine gearboxes.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"47 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"132231000","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Tonal noise is one of the main sources of noise generated by cooling fan module in automobile. The methods to decrease the tonal noise become more and more important with the popularity of automobiles and the demand for comfort increasing. Finding a quick way to predict the tonal noise in the early design stage will help the designer shorten the design circle and improve the efficiency. The traditional way, uneven spacing rotor blades, to modulate tonal components in noise spectrum by controlling factor “A” show positive influence in harmonic 2 and 3 (H2 and H3) but cannot work for harmonic 1 (H1) subjects to limitation of blade overlap as shown in the test result. Further step to modulate the stator vanes by adjusting parameter “χ” can attenuate the H1 dramatically based on “Venetian blind” theory of stator vanes. At last, the new cooling fan module design method to control tonal noise in design stage by “A” in rotor blade design and “χ” in stator vanes is reviewed and implemented in two cases of 7 rotor-blades & 6 stator-vanes and 7 rotor-blades & 8 stator-vanes. The experimental results agree well with the theoretical prediction.
{"title":"Tonal Noise Control of Cooling Fan Module by Using Modulation Principles on Both Rotor and Stator","authors":"Zhigang Peng, H. Ouyang, Yadong Wu, J. Tian","doi":"10.1115/GT2018-76300","DOIUrl":"https://doi.org/10.1115/GT2018-76300","url":null,"abstract":"Tonal noise is one of the main sources of noise generated by cooling fan module in automobile. The methods to decrease the tonal noise become more and more important with the popularity of automobiles and the demand for comfort increasing. Finding a quick way to predict the tonal noise in the early design stage will help the designer shorten the design circle and improve the efficiency. The traditional way, uneven spacing rotor blades, to modulate tonal components in noise spectrum by controlling factor “A” show positive influence in harmonic 2 and 3 (H2 and H3) but cannot work for harmonic 1 (H1) subjects to limitation of blade overlap as shown in the test result. Further step to modulate the stator vanes by adjusting parameter “χ” can attenuate the H1 dramatically based on “Venetian blind” theory of stator vanes. At last, the new cooling fan module design method to control tonal noise in design stage by “A” in rotor blade design and “χ” in stator vanes is reviewed and implemented in two cases of 7 rotor-blades & 6 stator-vanes and 7 rotor-blades & 8 stator-vanes. The experimental results agree well with the theoretical prediction.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"34 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133144617","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
With the aim of analysing the efficiency of leading edge serrations under realistic conditions, an experimental rig was developed where a ducted low-speed fan is installed that allows to gather data of both, aerodynamic and aeroacoustic nature. Turbulent inflow conditions were generated via biplane-square grids, resulting in turbulence intensities of different magnitude and of high isotropic character that were quantified by use of hotwire measurements. The fan blades were designed according to the NACA65(12)-10 profile with interchangeable features and an independently adjustable angle of attack. Altogether, five different parameters can be analysed, namely the serration amplitude and wavelength, the angle of attack, the inflow turbulence and the rotational speed. In addition, the blade design allows for a variation of the blade skew, sweep and dihedral as well. The presented work focusses on validating and optimising the test rig as well as a detailed quantification of the turbulent inflow conditions. Furthermore, first aerodynamic and aeroacoustic results of fan blades with straight leading edges are compared to those of serrated leading edges. The aerodynamic performance was found to be mainly affected by the serrations as a function of the serration amplitude. Aeroacoustically, a clear sensitivity towards different incoming turbulence intensities and serration parameters was detected, showing significant broadband noise reduction below 2 kHz with an overall noise reduction of ΔOASPL = 3.4 dB at maximum serration amplitudes and minimum wavelengths.
{"title":"Optimised Test Rig for Measurements of Aerodynamic and Aeroacoustic Performance of Leading Edge Serrations in Low-Speed Fan Application","authors":"T. Biedermann, F. Kameier, C. Paschereit","doi":"10.1115/GT2018-75369","DOIUrl":"https://doi.org/10.1115/GT2018-75369","url":null,"abstract":"With the aim of analysing the efficiency of leading edge serrations under realistic conditions, an experimental rig was developed where a ducted low-speed fan is installed that allows to gather data of both, aerodynamic and aeroacoustic nature. Turbulent inflow conditions were generated via biplane-square grids, resulting in turbulence intensities of different magnitude and of high isotropic character that were quantified by use of hotwire measurements. The fan blades were designed according to the NACA65(12)-10 profile with interchangeable features and an independently adjustable angle of attack. Altogether, five different parameters can be analysed, namely the serration amplitude and wavelength, the angle of attack, the inflow turbulence and the rotational speed. In addition, the blade design allows for a variation of the blade skew, sweep and dihedral as well. The presented work focusses on validating and optimising the test rig as well as a detailed quantification of the turbulent inflow conditions. Furthermore, first aerodynamic and aeroacoustic results of fan blades with straight leading edges are compared to those of serrated leading edges. The aerodynamic performance was found to be mainly affected by the serrations as a function of the serration amplitude. Aeroacoustically, a clear sensitivity towards different incoming turbulence intensities and serration parameters was detected, showing significant broadband noise reduction below 2 kHz with an overall noise reduction of ΔOASPL = 3.4 dB at maximum serration amplitudes and minimum wavelengths.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"38 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114802776","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper will discuss the development and testing of an electric drive option designed for the propulsion system of the US Navy’s DDG 51 Class ships. It will briefly explain the history of the Hybrid Electric Drive (HED) program, including that of its predecessor, Proof of Concept (PoC), and the HED’s planned shipboard installation schedule. Operating at lower ship speeds, in a range where the currently installed propulsion gas turbines are less fuel efficient, the HED is expected to increase the ship’s fuel economy, allowing the ship to remain on station accomplishing its mission for a longer period of time. This paper will discuss how the gas turbine propulsion system, in concert with the HED, will be used to provide the most fuel efficient drive combination for various operating scenarios. Also covered will be a description of the major stakeholders involved in the HED’s development and implementation along with some of the constraints and challenges that were encountered in the testing phase of the program, both at the OEM facilities and at the US Navy’s Land Based Engineering Site (LBES) in Philadelphia PA. Planned fuel economy testing results obtained at the LBES facility will also be presented, intended to determine an estimate of the fuel savings that can be expected when the system is first placed in service on USS TRUXTUN (DDG 103) July 2018.
{"title":"Development and Testing of the Hybrid Electric Drive Program for the Navy’s DDG 51 Class Ships","authors":"Gianfranco P. Buonamici, Michaela Schauble","doi":"10.1115/GT2018-76928","DOIUrl":"https://doi.org/10.1115/GT2018-76928","url":null,"abstract":"This paper will discuss the development and testing of an electric drive option designed for the propulsion system of the US Navy’s DDG 51 Class ships. It will briefly explain the history of the Hybrid Electric Drive (HED) program, including that of its predecessor, Proof of Concept (PoC), and the HED’s planned shipboard installation schedule.\u0000 Operating at lower ship speeds, in a range where the currently installed propulsion gas turbines are less fuel efficient, the HED is expected to increase the ship’s fuel economy, allowing the ship to remain on station accomplishing its mission for a longer period of time. This paper will discuss how the gas turbine propulsion system, in concert with the HED, will be used to provide the most fuel efficient drive combination for various operating scenarios. Also covered will be a description of the major stakeholders involved in the HED’s development and implementation along with some of the constraints and challenges that were encountered in the testing phase of the program, both at the OEM facilities and at the US Navy’s Land Based Engineering Site (LBES) in Philadelphia PA.\u0000 Planned fuel economy testing results obtained at the LBES facility will also be presented, intended to determine an estimate of the fuel savings that can be expected when the system is first placed in service on USS TRUXTUN (DDG 103) July 2018.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"41 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123226196","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Electric propulsion is being considered for a wide range of airframes from large commercial transports to the small Unmanned Aerial Systems (UASs). These electric systems, especially for small fixed wing UASs and quadcopters, need to be both efficient and quiet if they are to operate in an urban/populated environment or used in an Intelligence, Surveillance, and Reconnaissance (ISR) scenario. A propeller test facility was developed to record propeller performance and sound generation in the near field behind UAS propellers. The question of defining near and far field noise was studied by characterizing sound decay with distance from a UAS propeller. Defining near and far field noise is a subject that is not addressed well in the literature. Far field noise generally follows the 1/r decay rate and near field does not. Behind the propeller there are other flow field interactions that also change the decay rate, which this study illustrates. The data presented in this paper shows the difficulty in measuring sound around a UAS propeller and begins to resolve this topic. Previous UAS propeller design work by the authors resulted in propellers that were quieter in the near field and at the same time more efficient. Their studies showed RPM and tip vortex formation both contribute significantly to propeller sound generation. Disrupting the tip vortex formation should decrease the noise being generated. The current work extends these initial findings and examines the noise generation of a stock quadcopter propeller from a DJI Phantom 2 platform. One inch aft of the plane of rotation, this propeller, a 9.4 × 5.0, has a peak sound pressure level (SPL) of approximately 118 dBA under normal static operation producing 0.7 lbf of thrust at approximately 5900 RPM. Modifications were made to four stock propellers by cutting a notch perpendicular to the leading edge of the propeller at the 0.75 r/R and 0.87 r/R locations. The notches were of different depths and widths. Of the modifications, three of the configurations did not noticeably decrease the sound. However; the final configuration reduced the peak near field SPL to 111 dBA, a 6% reduction in dBA over the stock configuration corresponding to a greater than 50% reduction in sound generation. Smoke visualization confirms that a notch located at 0.87 r/R effectively disrupts the tip vortex formation, causing the tip vortices to dissipate much earlier than the stock propeller without the notch. Examining the noise frequency spectrums associated with both the stock and the modified propeller also confirm that the notch changes the magnitude and frequency distribution of the sound being generated.
{"title":"Near and Far Field Noise Decay From a Quadcopter Propeller With and Without a Leading Edge Notch","authors":"K. V. Treuren, C. Wisniewski, E. Cinnamon","doi":"10.1115/GT2018-75973","DOIUrl":"https://doi.org/10.1115/GT2018-75973","url":null,"abstract":"Electric propulsion is being considered for a wide range of airframes from large commercial transports to the small Unmanned Aerial Systems (UASs). These electric systems, especially for small fixed wing UASs and quadcopters, need to be both efficient and quiet if they are to operate in an urban/populated environment or used in an Intelligence, Surveillance, and Reconnaissance (ISR) scenario.\u0000 A propeller test facility was developed to record propeller performance and sound generation in the near field behind UAS propellers. The question of defining near and far field noise was studied by characterizing sound decay with distance from a UAS propeller. Defining near and far field noise is a subject that is not addressed well in the literature. Far field noise generally follows the 1/r decay rate and near field does not. Behind the propeller there are other flow field interactions that also change the decay rate, which this study illustrates. The data presented in this paper shows the difficulty in measuring sound around a UAS propeller and begins to resolve this topic. Previous UAS propeller design work by the authors resulted in propellers that were quieter in the near field and at the same time more efficient. Their studies showed RPM and tip vortex formation both contribute significantly to propeller sound generation. Disrupting the tip vortex formation should decrease the noise being generated. The current work extends these initial findings and examines the noise generation of a stock quadcopter propeller from a DJI Phantom 2 platform. One inch aft of the plane of rotation, this propeller, a 9.4 × 5.0, has a peak sound pressure level (SPL) of approximately 118 dBA under normal static operation producing 0.7 lbf of thrust at approximately 5900 RPM. Modifications were made to four stock propellers by cutting a notch perpendicular to the leading edge of the propeller at the 0.75 r/R and 0.87 r/R locations. The notches were of different depths and widths. Of the modifications, three of the configurations did not noticeably decrease the sound. However; the final configuration reduced the peak near field SPL to 111 dBA, a 6% reduction in dBA over the stock configuration corresponding to a greater than 50% reduction in sound generation. Smoke visualization confirms that a notch located at 0.87 r/R effectively disrupts the tip vortex formation, causing the tip vortices to dissipate much earlier than the stock propeller without the notch. Examining the noise frequency spectrums associated with both the stock and the modified propeller also confirm that the notch changes the magnitude and frequency distribution of the sound being generated.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"104 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131519582","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Under-race lubrication is the main method for the main shaft bearing of aero-engine which with higher performance. Oil scoop is an important part of the under-race lubrication structure, which plays an important role in capturing oil coming out of a stationary jet nozzle, and the efficiency of oil capture has great influence on the performance of the under-race lubrication. In this paper, a reasonable numerical simulation method is used to calculate a certain radial oil scoop. The velocity distribution of the internal air field in the lubrication structure and the oil distribution of the oil-gas two phase flow field were calculated and the scoop efficiency under different working conditions were calculated. The scoop efficiency under the three oil jet nozzles was verified by the test data. Finally, the influence of the shaft rotation speed, the oil flow rate and the number of the oil nozzles on the scoop efficiency of the radial scoop is analyzed, and the reason of these regularities is analyzed in detail. The result of this study may provide an idea or method for the optimization and improvement of oil scoop with similar structure.
{"title":"Simulation and Analysis of Oil Scoop Capture Efficiency","authors":"Lyu Yaguo, Jiang Le, Liu Zhenxia, H. Jianping","doi":"10.1115/GT2018-75989","DOIUrl":"https://doi.org/10.1115/GT2018-75989","url":null,"abstract":"Under-race lubrication is the main method for the main shaft bearing of aero-engine which with higher performance. Oil scoop is an important part of the under-race lubrication structure, which plays an important role in capturing oil coming out of a stationary jet nozzle, and the efficiency of oil capture has great influence on the performance of the under-race lubrication. In this paper, a reasonable numerical simulation method is used to calculate a certain radial oil scoop. The velocity distribution of the internal air field in the lubrication structure and the oil distribution of the oil-gas two phase flow field were calculated and the scoop efficiency under different working conditions were calculated. The scoop efficiency under the three oil jet nozzles was verified by the test data. Finally, the influence of the shaft rotation speed, the oil flow rate and the number of the oil nozzles on the scoop efficiency of the radial scoop is analyzed, and the reason of these regularities is analyzed in detail. The result of this study may provide an idea or method for the optimization and improvement of oil scoop with similar structure.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"50 2","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"120850517","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Igor Neifach, Gi-Don Na, F. Kameier, Nils Springer, M. Wichers
This paper deals with the reduction of aerodynamically generated noise in passenger car Cooling-Fan-Modules (CFM), caused by the interaction between the impeller and the downstream-located strut configuration of the axial fan. Even after the car engine is switched off, the fan remains active, as long as cooling is required for certain vehicle components. Especially after a car has been parked in closed parking areas, in close proximity to residential buildings or public places, the noise emission can be a problem. This issue is addressed by dampening the rotor-stator-interaction through passive construction measures. In order to ensure optimal noise reduction, 8 critical design features of the struts are identified and investigated using statistical design of experiment methods (DoE). Based on the results, dedicated insights about the effects of concrete strut features on significant regions of the acoustic fan spectrum are obtained. Furthermore, an optimized strut configuration is derived and metrologically validated using a polyoptimization method. Compared to a current serial baseline configuration, a reduction of the overall sound pressure level by 2.6 dB(A), as well as a reduction of the blade passage frequency tone by 17.6 dB(A) is achieved.
{"title":"Aeroacoustic Optimization of a Pressure-Side Strut Configuration for Subsonic Axial Fans Using Statistical-Empirical Modelling","authors":"Igor Neifach, Gi-Don Na, F. Kameier, Nils Springer, M. Wichers","doi":"10.1115/GT2018-76348","DOIUrl":"https://doi.org/10.1115/GT2018-76348","url":null,"abstract":"This paper deals with the reduction of aerodynamically generated noise in passenger car Cooling-Fan-Modules (CFM), caused by the interaction between the impeller and the downstream-located strut configuration of the axial fan. Even after the car engine is switched off, the fan remains active, as long as cooling is required for certain vehicle components. Especially after a car has been parked in closed parking areas, in close proximity to residential buildings or public places, the noise emission can be a problem. This issue is addressed by dampening the rotor-stator-interaction through passive construction measures. In order to ensure optimal noise reduction, 8 critical design features of the struts are identified and investigated using statistical design of experiment methods (DoE). Based on the results, dedicated insights about the effects of concrete strut features on significant regions of the acoustic fan spectrum are obtained. Furthermore, an optimized strut configuration is derived and metrologically validated using a polyoptimization method. Compared to a current serial baseline configuration, a reduction of the overall sound pressure level by 2.6 dB(A), as well as a reduction of the blade passage frequency tone by 17.6 dB(A) is achieved.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"45 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115986705","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Uncertainties surrounding the influence of Reynolds number on the performance of air handling turbomachines are as old as the study of turbomachinery fluid dynamics. In particular, all low-speed turbomachines and most axial-flow fans feature Reynolds numbers that are often lower than the critical value, above which the literature states a limited dependency of blades cascade aerodynamics on Reynolds number. Testing standards already account for this well-known issue, which arises mainly in the case of geometrically similar fans of different size and/or operating conditions. On the other hand, one of the main practical issues in the design of low-speed machines is the disagreement among the most authoritative sources on the value of the critical Reynolds number for axial fans. The many definitions of Reynolds number, which are suited to either fan design purposes or fan performance assessment, introduce additional problems, as the corresponding values may differ by orders of magnitude depending on the chosen definition. A less debated issue deals with the effect of Reynolds number on global performance and efficiency parameters for different axial-flow fan configurations. This paper reports pressure and efficiency data measured at several rotational speeds of four axial fans that feature different configurations, hub-to-tip ratios, sizes and surface finishes. In particular, the tests consider two 315mm and one 630mm tube-axial fans, and one 800mm vane-axial fan with preswirler blading. Data on two vane-axial fans with straightener, and one preswirler-rotor-stator stage, available in the literature, widen the discussion on the Reynolds number effect on the entire category of single-stage axial fans.
{"title":"Experimental Investigation of the Effect of Reynolds Number on the Efficiency of Single-Stage Axial Fans","authors":"M. Masi, S. Castegnaro, A. Lazzaretto","doi":"10.1115/GT2018-76909","DOIUrl":"https://doi.org/10.1115/GT2018-76909","url":null,"abstract":"Uncertainties surrounding the influence of Reynolds number on the performance of air handling turbomachines are as old as the study of turbomachinery fluid dynamics. In particular, all low-speed turbomachines and most axial-flow fans feature Reynolds numbers that are often lower than the critical value, above which the literature states a limited dependency of blades cascade aerodynamics on Reynolds number. Testing standards already account for this well-known issue, which arises mainly in the case of geometrically similar fans of different size and/or operating conditions.\u0000 On the other hand, one of the main practical issues in the design of low-speed machines is the disagreement among the most authoritative sources on the value of the critical Reynolds number for axial fans. The many definitions of Reynolds number, which are suited to either fan design purposes or fan performance assessment, introduce additional problems, as the corresponding values may differ by orders of magnitude depending on the chosen definition.\u0000 A less debated issue deals with the effect of Reynolds number on global performance and efficiency parameters for different axial-flow fan configurations.\u0000 This paper reports pressure and efficiency data measured at several rotational speeds of four axial fans that feature different configurations, hub-to-tip ratios, sizes and surface finishes. In particular, the tests consider two 315mm and one 630mm tube-axial fans, and one 800mm vane-axial fan with preswirler blading. Data on two vane-axial fans with straightener, and one preswirler-rotor-stator stage, available in the literature, widen the discussion on the Reynolds number effect on the entire category of single-stage axial fans.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"193 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121615662","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
NASA’s Environmentally Responsible Aviation (ERA) project supported the collection of a large stereoscopic Particle Image Velocimetry (PIV) dataset of the Open Rotor Propulsion Rig (ORPR) in the 9 × 15 Low Speed Wind Tunnel at the NASA Glenn Research Center (GRC). The data collection effort acquired a volume of three component velocity measurements composed of 30 planes from near the hub radially outward towards the tip. The PIV cameras and laser were mounted to a traverse that allowed the entire data acquisition system to move from plane to plane. The PIV data acquisition was triggered on the front rotor such that the front rotor was always in the same position for each acquisition event. The aft rotor position was not recorded and varied randomly during the acquisition. Because the position of the aft rotor was not synchronized to either the forward rotor or the camera it was necessary to separate individual PIV images based off of the phase of the aft rotor before they could be processed. The phase of the aft rotor was determined by locating the outline of the rotor in the PIV images and determining its position relative to a known point. This process was conducted by an image processing algorithm. Previous algorithms were able to make a relatively accurate 3D model of the wake between the forward and aft rotors, however some small inaccuracies were present. Improvements to this algorithm allowed for more accurate phase averaging, which yielded an improved PIV dataset. Both rotors were set to the same nominal rotational speed, however variations in motor control and other physical mechanisms allowed for some differences in the true RPM of the two rotors. The effects of the aft rotor on the front rotor blade wakes within the inter-rotor flow field of the ORPR were examined. The aft rotor potential field was shown to have significant upstream impact on the front rotor wakes, altering their topology approaching the aft rotor. The wake strength was quantified through determination of the minimum velocities that occurred within a wake while a defined velocity deficit from the free stream velocity was used to determine the width of the wake. This work has potential applications to the radiated acoustics and efficiency of open rotor propulsion systems.
{"title":"Effect of Aft Rotor on Forward Rotor Blade Wakes in Open Rotor Propulsion Systems","authors":"C. DeLorenzo, J. Deleon, Paul E. Slaboch","doi":"10.1115/GT2018-76183","DOIUrl":"https://doi.org/10.1115/GT2018-76183","url":null,"abstract":"NASA’s Environmentally Responsible Aviation (ERA) project supported the collection of a large stereoscopic Particle Image Velocimetry (PIV) dataset of the Open Rotor Propulsion Rig (ORPR) in the 9 × 15 Low Speed Wind Tunnel at the NASA Glenn Research Center (GRC). The data collection effort acquired a volume of three component velocity measurements composed of 30 planes from near the hub radially outward towards the tip. The PIV cameras and laser were mounted to a traverse that allowed the entire data acquisition system to move from plane to plane. The PIV data acquisition was triggered on the front rotor such that the front rotor was always in the same position for each acquisition event. The aft rotor position was not recorded and varied randomly during the acquisition.\u0000 Because the position of the aft rotor was not synchronized to either the forward rotor or the camera it was necessary to separate individual PIV images based off of the phase of the aft rotor before they could be processed. The phase of the aft rotor was determined by locating the outline of the rotor in the PIV images and determining its position relative to a known point. This process was conducted by an image processing algorithm. Previous algorithms were able to make a relatively accurate 3D model of the wake between the forward and aft rotors, however some small inaccuracies were present. Improvements to this algorithm allowed for more accurate phase averaging, which yielded an improved PIV dataset. Both rotors were set to the same nominal rotational speed, however variations in motor control and other physical mechanisms allowed for some differences in the true RPM of the two rotors.\u0000 The effects of the aft rotor on the front rotor blade wakes within the inter-rotor flow field of the ORPR were examined. The aft rotor potential field was shown to have significant upstream impact on the front rotor wakes, altering their topology approaching the aft rotor. The wake strength was quantified through determination of the minimum velocities that occurred within a wake while a defined velocity deficit from the free stream velocity was used to determine the width of the wake. This work has potential applications to the radiated acoustics and efficiency of open rotor propulsion systems.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"108 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114227375","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
As a prospective pressure gain combustion technology, detonation combustion has obvious potential for greatly increasing the thermodynamic performance of marine gas turbine due to its advantage in low entropy generation, fast heat release and self-pressurization. In this paper, a thermodynamic cycle model of detonation combustion based marine gas turbine is established considering the variable specific heat capacity. On this basis, a comparative analysis is investigated to discuss the effects of different factors on the performance enhancement of marine gas turbine by using detonation combustion. The results demonstrate that compared to the conventional deflagration combustion, detonation combustion can significantly improve the thermodynamic performance of marine gas turbine under various condition. As far as the present study is concerned, the thermal cycle efficiency can be increased to 42.97∼46.76%. Besides, it is found that the effects of pressure ratio on performance enhancements of marine gas turbine are higher than those of atmospheric temperature and temperature ratio. When pressure ratio is ranged from 13 to 30, both thermal cycle efficiency and specific power enhancements are about 20∼27%.
{"title":"Thermodynamic Performance Enhancement of Marine Gas Turbine by Using Detonation Combustion","authors":"Ningbo Zhao, Hongtao Zheng, Xueyou Wen, Dongming Xiao","doi":"10.1115/GT2018-75493","DOIUrl":"https://doi.org/10.1115/GT2018-75493","url":null,"abstract":"As a prospective pressure gain combustion technology, detonation combustion has obvious potential for greatly increasing the thermodynamic performance of marine gas turbine due to its advantage in low entropy generation, fast heat release and self-pressurization. In this paper, a thermodynamic cycle model of detonation combustion based marine gas turbine is established considering the variable specific heat capacity. On this basis, a comparative analysis is investigated to discuss the effects of different factors on the performance enhancement of marine gas turbine by using detonation combustion. The results demonstrate that compared to the conventional deflagration combustion, detonation combustion can significantly improve the thermodynamic performance of marine gas turbine under various condition. As far as the present study is concerned, the thermal cycle efficiency can be increased to 42.97∼46.76%. Besides, it is found that the effects of pressure ratio on performance enhancements of marine gas turbine are higher than those of atmospheric temperature and temperature ratio. When pressure ratio is ranged from 13 to 30, both thermal cycle efficiency and specific power enhancements are about 20∼27%.","PeriodicalId":114672,"journal":{"name":"Volume 1: Aircraft Engine; Fans and Blowers; Marine","volume":"148 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2018-06-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114635123","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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