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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1978

Reduction of fan noise in an anechoic chamber by reducing chamber wall induced inlet flow disturbances

The difference between the flight and ground static noise of turbofan engines presents a significant problem in engine noise testing. The additional noise for static testing has been attributed to inlet flow disturbances or turbulence interacting with the fan rotor. In an attempt to determine a possible source of inflow disturbances entering fans tested in the Lewis Research Center anechoic chamber, the inflow field was studied using potential flow analysis. These potential flow calculations indicated that there was substantial flow over the wall directly behind the fan inlet that could produce significant inflow disturbances. Fan noise tests were run with various extensions added to the fan inlet to move the inlet away from this backwall and thereby reduce the inlet flow disturbances. Significant noise reductions were observed with increased inlet length. Over 5 db reduction of the blade passage tone sound power level was observed between the shortest and longest inlets at 90% fan speed and the first overtone was reduced 9 db. High frequency broadband noise was also reduced.

Dittmar, J. H.↗
NASA NTRS1978

Reduction of fan noise in an anechoic chamber by reducing chamber wall induced inlet flow disturbances

The difference between the flight and ground static noise of turbofan engines has been identified as a significant problem in engine noise testing. The additional noise for static testing has been attributed to inlet flow disturbances or turbulence interacting with the fan rotor. In an attempt to determine a possible source of inflow disturbances entering fans tested in the Lewis Research Center anechoic chamber the inflow field was studied using potential flow analysis. These potential flow calculations indicated that there was substantial flow over the wall directly behind the fan inlet that could produce significant inflow disturbances. Fan noise tests were run with various extensions added to the fan inlet to move the inlet away from this backwall and thereby reduce the inlet flow disturbances. Significant noise reductions were observed with increased inlet length. Over 5 dB reduction of the blade passage tone sound power level was observed between the shortest and longest inlets at 90% fan speed and the first overtone was reduced 9 dB. High frequency broadband noise was also reduced.

Dittmar, J. H.↗
NASA NTRS1975

Demonstration of short-haul aircraft aft noise reduction techniques on a twenty inch (50.8) diameter fan, volume 2

Aft fan noise reduction techniques were investigated. The 1/3 octave band sound data were plotted with the following plots included: perceived noise level vs acoustic angle at 2 fan speeds; PWL vs frequency at 2 fan speeds; and sound pressure level vs frequency at 2 aft angles and 2 fan speeds. The source noise plots included: band pass filter sound pressure level vs acoustic angle at 2 fan speeds; and 2nd harmonic SPL acoustic angle at 2 fan speeds.

Stimpert, D. L.↗
NASA NTRS1981

Static test of a fan-powered chin nozzle for V/STOl applications

The performance of a "chin" nozzle which diverts flow in a downward direction immediately downstream of a fan typical of designs suitable for V/STOL A applications was evaluated. Back pressure distortion to the fan and fan discharge pressure distortion were also measured. Results show that the distortion is significant at the closest spacing between the fan exit and cascade entrance tested, and that the chin nozzle performance deteriorates with increased flow diversion to the chin nozzle. Color oil flow visualization on video tape and still photos were also obtained. Tests were conducted behind a 12" model fan in the NASA-Lewis fan calibration facility.

Salemann, V.↗
NASA NTRS1983

Aerodynamic performance of a fan stage utilizing Variable Inlet Guide Vanes (VIGVs) for thrust modulation

An experimental research program was conducted in the Lewis Research Center's 9x15-foot (2.74x4.57 m) low speed wind tunnel to evaluate the aerodynamic performance of an inlet and fan system with variable inlet guide vanes (VIGVs) for use on a subsonic V/STOL aircraft. At high VIGV blade angles (lower weight flow and thrust levels), the fan stage was stalled over a major portion of its radius. In spite of the stall, fan blade stresses only exceeded the limits at the most extreme flow conditions. It was found that inlet flow separation does not necessarily lead to poor inlet performance or adverse fan operating conditions. Generally speaking, separated inlet flow did not adversely affect the fan blade stress levels. There were some cases, however, at high VIGV angles and high inlet angles-of-attack where excessive blade stress levels were encountered. An evaluation term made up of the product of the distortion parameter, K alpha, the weight flow and the fan pressure ratio minus one, was found to correlate quite well with the observed blade stress results.

Woollett, R. R.↗
NASA NTRS1983

The aerodynamic design and performance of the General Electric/NASA EEE fan

The aerodynamic design and test results of the fan and quarter-stage component for the GE/NASA Energy Efficient Engine (EEE) are presented. The fan is a high bypass ratio, single-stage design having 32 part-span shrouded rotor blades, coupled with a unique quarter-stage arrangement that provides additional core-stream pressure ratio and particle separation. The fan produces a bypass pressure ratio of 1.65 at the exit of the low aspect ratio vane/frame and a core-stream pressure ratio of 1.67 at the entrance to the core frame struts. The full-scale fan vehicle was instrumented, assembled and tested as a component in November 1981. Performance mapping was conducted over a range of speeds and bypass ratios using individually-controlled bypass and core-stream discharge valves. The fan bypass and core-stream test data showed excellent results, with the fan exceeding all performance goals at the important engine operating conditions.

Sullivan, T. J.↗
NASA NTRS1983

Aerodynamic performance of a fan stage utilizing variable inlet guide vanes (VIGV's) for thrust modulation

An experimental research program was conducted in the Lewis Research Center's 9 x 15-foot (2.74 x 4.57 m) low speed wind tunnel to evaluate the aerodynamic performance of an inlet and fan system with variable inlet guide vanes (VIGVs) for use on a subsonic V/STOL aircraft. At high VIGV blade angles (lower weight flow and thrust levels), the fan stage was stalled over a major portion of its radius. In spite of the stall, fan blade stresses only exceeded the limits at the most extreme flow conditions. It was found that inlet flow separation does not necessarily lead to poor inlet performance or adverse fan operating conditions. Generally speaking, separated inlet flow did not adversely affect the fan blade stress levels. There were some cases, however, at high VIGV angles and high inlet angles-of-attack where excessive blade stress levels were encountered. An evaluation term made up of the product of the distortion parameter, K alpha, the weight flow and the fan pressure ratio minus one, was found to correlate quite well with the observed blade stress results. Previously announced in STAR as N83-27957

Woollett, R. R.↗
NASA NTRS1983

Energy efficient engine. Fan and quarter-stage component performance report

The fan configuration for the general Electric/NASA Energy Efficient Engine was selected following an extensive preliminary design study. The fan has an inlet radius ratio of 0.342 and a specific flowrate of 208.9 Kg/sec/sq. m (42.8 1bm/sec/sq. ft). The design corrected tip speed is 411.5 m/sec (1350 ft/sec) producing a bypass flow total-pressure ratio of 1.65 and a core flow total-pressure ratio of 1.6. The design bypass ratio is 6.8. The aerodynamic design point corresponds to the maximum climb power setting at Mach 0.8 and 10.67 Km (35,000 ft) altitude. The fully-instrumented fan component was tested in the Lynn Large Fan Test Facility in 1981. The overall performance results, reported herein, showed excellent fan performance with the fan meeting all of its component test goals of flow, efficiency and stall margin.

Cline, S. J.↗
NASA NTRS1989

A supersonic through-flow fan engine airframe integration study

A study is undertaken to investigate the engine airframe integration effects for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic through-flow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic through-flow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic through-flow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum lift-to-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from 1.10 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration.

Barnhart, Paul J.↗
NASA NTRS1989

A supersonic through-flow fan engine airframe integration study

Engine airframe integration effects are investigated for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic through-flow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic through-flow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic through-flow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum lift-to-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from 1.10 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration.

Barnhart, Paul J.↗
NASA NTRS1992

A concept for a counterrotating fan with reduced tone noise

As subsonic jet engine designs incorporate higher bypass ratios to reduce jet noise and increase engine cycle efficiency, the fan noise becomes a significant part of the perceived total noise. The conventional method of reducing fan tone noise is to design a low tip-speed device. An alternative approach of using a counterrotating fan with a high number of rotor blades is investigated in this report. The source of noise at the blade passing frequency of this device is the rotor-only mechanism, which is cut off for a subsonic tip speed rotor. The interaction noise occurs at twice the blade passing frequency, which, for this fan, was shifted high enough in frequency to be above the perceived noise rating range. The result was a counterrotating fan which had more potential for tone noise reduction than does the conventional fan. A potential broadband noise reduction was also indicated.

Dittmar, James H.↗
NASA NTRS1993

Noise Trends of a 0.5 M (20 In.) Diameter Supersonic Throughflow Fan as Measured in an Unmodified Compressor Aerodynamic Test Facility

The tone noise levels of a supersonic throughflow fan were measured at subsonic and supersonic axial duct Mach numbers. The noise in the inlet plenum showed no blade passing and harmonic tones at subsonic or supersonic axial flow conditions. At subsonic axial flow conditions, the supersonic throughflow fan showed no inlet plenum tones at fan operating conditions where tone noise had been previously measured for a subsonic fan design. This lower inlet-quadrant noise level for the supersonic throughflow fan was the result of high subsonic inlet velocities acting to reduce the noise propagating out the inlet. The fan noise, which was prevented from propagating upstream by the high subsonic inlet velocities, appeared to increase the noise in the exhaust duct at subsonic throughflow conditions. The exhaust duct noise decreased at supersonic axial throughflow Mach numbers, with the lowest blade passing and harmonic tones levels being observed at the design axial Mach number of 2.0. Multiple pure tone noise was observed in the inlet duct at subsonic axial flow Mach numbers but was seen only in the exhaust duct at supersonic axial flow conditions.

Dittmar, James H.↗
NASA NTRS1920

A Variable Speed Fan Dynamometer

Fan brakes used as absorption dynamometers in testing internal combustion engines have the disadvantage that a given fan will run only at one speed when the engine is delivering full power. In order to be able to vary the speed at which a given power will be absorbed, English manufacturers have for some time been using a cylindrical housing around the fan with one or two variable openings in the periphery. Here, results are given of tests conducted to determine how great a range of speed can be obtained from such a device. The tests show that a power ratio of five to 1 can be obtained, the power ratio being defined as the ratio of the power absorbed by the fan at a given speed with the outlet open to the power absorbed at the same speed with the second outlet closed. Data show that improvements in the design of the fan brake can make the speed ratio approach but not exceed a value of two to one. Also given here are a brief outline of previous work on fan brakes, a description of the experimental apparatus and methods used in the tests, and a more detailed statement of test results.

Wood, Karl D↗
NASA NTRS1942

Test of Single-Stage Axial-Flow Fan

A single-stage axial fan was built and tested in the shop of the propeller-research tunnel of the NACA. The fan comprised a simple 24-blade rotor having a diameter of 21 inches and a solidity of 0.86 and a set of 37 contravanes having a solidity of 1.33. The rotor was driven by a 25-horsepower motor capable of rotating at a speed of 3600 r.p.m. The fan was tested for volume, pressure, and efficiency over a range of delivery pressures and volumes for a wide range of contravane and blade-angle settings. The test results are presented in chart form in terms of nondimensional units in order that similar fans may be accurately designed with a minimum effort. The maximum efficiency (88 percent) was obtained by the fan at a blade angle of 30 degrees and a contravane angle of 70 degrees. An efficiency of 80 percent was obtained by the fan with the contravanes removed.

Bell, E Barton↗
NASA NTRS1993

Active control of fan-generated plane wave noise

Subsonic propulsion systems for future aircraft may incorporate ultra-high bypass ratio ducted fan engines whose dominant noise source is the fan with blade passage frequency less than 1000 Hz. This low frequency combines with the requirement of a short nacelle to diminish the effectiveness of passive duct liners. Active noise control is seen as a viable method to augment the conventional passive treatments. An experiment to control ducted fan noise using a time domain active adaptive system is reported. The control sound source consists of loudspeakers arrayed around the fan duct. The error sensor location is in the fan duct. The purpose of this experiment is to demonstrate that the in-duct error sensor reduces the mode spillover in the far field, thereby increasing the efficiency of the control system. In this first series of tests, the fan is configured so that predominantly zero order circumferential waves are generated. The control system is found to reduce the blade passage frequency tone significantly in the acoustic far field when the mode orders of the noise source and of the control source are the same. The noise reduction is not as great when the mode orders are not the same even though the noise source modes are evanescent, but the control system converges stably and global noise reduction is demonstrated in the far field. Further experimentation is planned in which the performance of the system will be evaluated when higher order radial and spinning modes are generated.

Gerhold, Carl H.↗
NASA NTRS1996

Active Control of Fan Noise: Feasibility Study: Theoretical Analysis for Coupling of Active Noise Control Actuator Ring Sources to an Annular Duct with Flow - Volume 6

The objective of this effort is to develop an analytical model for the coupling of active noise control (ANC) piston-type actuators that are mounted flush to the inner and outer walls of an annular duct to the modes in the duct generated by the actuator motion. The analysis will be used to couple the ANC actuators to the modal analysis propagation computer program for the annular duct, to predict the effects of active suppression of fan-generated engine noise sources. This combined program will then be available to assist in the design or evaluation of ANC systems in fan engine annular exhaust ducts. An analysis has been developed to predict the modes generated in an annular duct due to the coupling of flush-mounted ring actuators on the inner and outer walls of the duct. The analysis has been combined with a previous analysis for the coupling of modes to a cylindrical duct in a FORTRAN computer program to perform the computations. The method includes the effects of uniform mean flow in the duct. The program can be used for design or evaluation purposes for active noise control hardware for turbofan engines. Predictions for some sample cases modeled after the geometry of the NASA Lewis ANC Fan indicate very efficient coupling in both the inlet and exhaust ducts for the m = 6 spinning mode at frequencies where only a single radial mode is cut-on. Radial mode content in higher order cut-off modes at the source plane and the required actuator displacement amplitude to achieve 110 dB SPL levels in the desired mode were predicted. Equivalent cases with and without flow were examined for the cylindrical and annular geometry, and little difference was found for a duct flow Mach number of 0.1. The actuator ring coupling program will be adapted as a subroutine to the cylindrical duct modal analysis and the exhaust duct modal analysis. This will allow the fan source to be defined in terms of characteristic modes at the fan source plane and predict the propagation to the arbitrarily-located ANC source plane. The actuator velocities can then be determined to generate the anti-phase mode. The resulting combined fan source/ANC pressure can then be calculated at any desired wall sensor position. The actuator velocities can be determined manually or using a simulation of a control system feedback loop. This will provide a very useful ANC system design and evaluation tool.

Kraft, R. E.↗
NASA NTRS2000

A Comparison of Measured Tone Modes for Two Low Noise Propulsion Fans

The acoustic modes for two low tip speed propulsion fans were measured to examine the effects of fan tip speed, at constant pressure ratio. A continuously rotating microphone method was used that provided the complete modal structure (circumferential and radial order) at the fundamental and second harmonic of the blade passing tone as well as most of the third harmonic modes. The fans are compared in terms of their rotor/stator interaction modal power, and total tone power. It was hoped that the lower tip speed might produce less noise. This was not the case. The higher tip speed fan, at both takeoff and cutback speeds, had lower tone and interaction levels. This could be an indication that the higher aerodynamic loading required to produce the same pressure ratio for the lower tip speed fan resulted in a greater velocity deficit in the blade wakes and thus more noise. Results consistent with expected rotor transmission effects were noted in the inlet modal structures of both fans.

Heidelberg, Laurence J.↗
NASA NTRS2003

A Fan Design that Meets the NASA Aeronautics Noise Goals

A fan concept was previously identified that would meet the NASA aeronautics goal of a 20 EPNdB reduction in aircraft noise. This was a 2-stage fan with a pressure ratio of 1.15 and a 460 ft/sec tip speed. The 2 stages were identical so that, with the proper synchrophasing, noise from one stage could be used to cancel noise from the other stage. This paper documents the aerodynamic design of the 2-stage fan concept in a 22-in. diameter size for testing in the NASA Glenn 9- by 15-ft wind tunnel. A set of rotor and stator coordinates are listed in the report. Stress and flutter analyses were done on these blades and showed that the design was structurally viable. A noise prediction code, using the blade coordinates and fan flows, indicated that the 2-stage fan would meet the goal of a 20 dB reduction in fan noise.

Dittmar, James↗