Static Engine Test Research Articles

Progress in prediction of jet noise and quantification of aircraft/engine noise components

The bypass ratio of newer turbofan engines has been increasing steadily and has reached ∼10; even higher bypass ratios are being considered for improving aircraft fuel efficiency. An accurate method for the prediction of jet noise over a wide range of bypass ratio, from ∼5 to ultra-high bypass ratios (∼20), is required to address current and future needs. The main objective of the current study is the development of a procedure for real-world application that would permit the accurate prediction of jet noise, which in turn would enable the quantification of the non-jet noise component. A new empirical method for prediction of noise from realistic dual-stream jets is developed and validated against an extensive database acquired at model scale; the range of validity covers a velocity ratio ( Vs/Vp) ≥ ∼0.6, as this is the range of interest in real turbofan engines. Accurate absolute spectral predictions at all angles and all cycle conditions are first demonstrated at model scale. The same method is then extended to full-scale predictions, both from static engine tests and airplane flyover tests. The range of Strouhal number of interest in full-scale tests spans ∼0.1 to ∼100. Nearfield effects on the jet noise from dual-stream jets have been quantified and a simple and improved procedure for incorporating spectral effects has been developed. Accurate spectral predictions are obtained for noise from static engine tests and flyover tests, over a wide range of engine operating conditions and bypass ratio. Therefore, the good quality and accuracy of the new prediction method have been confirmed at both model scale and full scale, with and without forward flight. An absolute prediction takes ∼1 s on a workstation, as only a single scaling equation is utilized. The application of the new prediction method in gaining better quantification and understanding of the non-jet noise components, and their reduction, is described.

Azimuthal Modes Measurement in an Intake Duct for a Turbofan Engine

The paper describes the acoustic mode measurements in a turbofan engine duct, performed with a sparse circumferential 100 - microphone array. The array was initially tested in the laboratory conditions in the anechoic chamber of Perm National Research Polytechnic University (PNRPU) with a mode synthesizer. It has been demonstrated that the measurement system correctly determines the synthesized modes in the duct. Then the array was installed on a turbofan engine, and the azimuthal structure of sound field in the engine intake duct was measured during static engine tests. The dominant modes were consistently determined for different engine runs albeit their amplitude changed from run to run, probably due to the absence of a turbulence control screen during the tests. As a result, it has been concluded that the measurement system can be used for measuring azimuthal structure of sound field in the ducts of new turbofan engines and thus can be instrumental in enhancement of the acoustic liner performance in turbofan engine noise reduction. A feasibility study on determining azimuthal modes in the engine duct with an array of microphones located in a limited angle range has been performed.

Infrasonic energy from orbital launch vehicles

Large, heavy-lift rockets have significant acoustic and infrasonic energy that can often be detected from a considerable distance. These sounds, under certain environmental conditions, can propagate hundreds of kilometers from the launch location. Thus, ground-based infrasound arrays can be used to monitor the low frequencies emitted by these large rocket launches. Multiple launches and static engine tests have been successfully recorded over many years using small infrasound arrays at various distances from the launch location. Infrasonic measurements using a 20 m array and parabolic equation modeling of a recent launch of an Aries V rocket at Wallops Island, Virginia, will be discussed.

Multiple pure tone noise prediction

This paper presents a fully numerical method for predicting multiple pure tones, also known as “Buzzsaw” noise. It consists of three steps that account for noise source generation, nonlinear acoustic propagation with hard as well as lined walls inside the nacelle, and linear acoustic propagation outside the engine. Noise generation is modeled by steady, part-annulus computational fluid dynamics (CFD) simulations. A linear superposition algorithm is used to construct full-annulus shock/pressure pattern just upstream of the fan from part-annulus CFD results. Nonlinear wave propagation is carried out inside the duct using a pseudo-two-dimensional solution of Burgers׳ equation. Scattering from nacelle lip as well as radiation to farfield is performed using the commercial solver ACTRAN/TM. The proposed prediction process is verified by comparing against full-annulus CFD simulations as well as against static engine test data for a typical high bypass ratio aircraft engine with hardwall as well as lined inlets. Comparisons are drawn against nacelle unsteady pressure transducer measurements at two axial locations as well as against near- and far-field microphone array measurements outside the duct.This is the first fully numerical approach (no experimental or empirical input is required) to predict multiple pure tone noise generation, in-duct propagation and far-field radiation. It uses measured blade coordinates to calculate MPT noise.

Psychoacoustic assessment of a new aircraft engine fan noise synthesis method

Simulation of aircraft flyover events can facilitate psychoacoustic studies exploring the effects of noise generated by future aircraft designs. The perceived realism of a simulated flyover event may be impacted by the perceived realism of the synthesized fan noise of the aircraft engine. Short-term fluctuations in tonal amplitude and frequency are important cues contributing to that perception of realism, but are not accounted for by predictions based on long-term averages. A new synthesis method has been developed at NASA Langley Research Center to generate realistic aircraft engine fan noise using predicted source noise directivities in combination with short-term fluctuations. In the new method, fluctuations in amplitude and frequency are included based upon analysis of static engine test data. Through psychoacoustic testing, this study assessed perceived effectiveness of the new synthesis method in generating realistic fan noise source. Realism was indirectly assessed by judging the similarity of synthesized sounds (with and without fluctuations) with recordings of fan noise. Results of ANOVA analyses indicated that subjects judged synthesized fan noise with fluctuations as being more similar to recordings than synthesized fan noise without fluctuations.

Hot Isostatic Pressing Technology for Defence and Space Applications

Hot isostatic pressing (HIP) technology has been established for the development of AISI-304 stainless steel and nickel base superalloy Inconel 718 integral turbine rotors, for liquid propulsion engine of Prithvi missile, and cryoengine of geostationary satellite launch vehicle (GSLV), respectively. Before making the full size rotors, the structure – property relationships in hot isostatic pressed (HIPed) 304 stainless steel and superalloy 718 were established. The HIPed steel and superalloy have shown near 100 per cent theoretical density, homogeneous, and fine grained microstructure. Their mechanical properties were found to be in agreement with those specified for the integral turbine rotors and hence, development of full size near net shaped integral turbine rotors was undertaken. The HIPed steel rotors subjected to the static engine tests have shown a satisfactory performance, and therefore a large number of rotors could be produced to fulfill the requirement of target labs. The HIP technology for the integral turbine rotors was found to be cost effective (about 50 per cent) over the conventional fabrication method which involves forging, machining, and welding of blades to the disk. The processing, structure, and properties of the HIPed 304 stainless steel and superalloy 718 in relation to the performance of integral turbine rotors for missile and space vehicle applications are discussed in this paper. Defence Science Journal, 2012, 62(1), pp.73 -80 , DOI:http://dx.doi.org/10.14429/dsj. 62.372

Shape Memory Alloy Based Morphing Aerostructures

In order to continue the current rate of improvements in aircraft performance, aircraft and components which are continuously optimized for all flight conditions, will be needed. Toward this goal morphing-capable, adaptive structures based on shape memory alloy (SMA) technology that enable component and system-level optimization at multiple flight conditions are being developed. This paper reviews five large-scale SMA based technology programs initiated by The Boeing Company. The SAMPSON smart inlet program showed that fully integrated SMA wire bundles could provide a fighter aircraft with a variable engine inlet capability. The reconfigurable rotor blade program demonstrated the ability of highly robust, controlled 55-Nitinol tube actuators to twist a rotor blade in a spin stand test to optimize rotor aerodynamic characteristics. The variable geometry chevron (VGC) program, which was the first use of 60-Nitinol for a major aerospace application, included a flight test and static engine test of the GE90-115B engine fitted with controlled morphing chevrons that reduced noise and increased engine efficiency. The deployable rotor tab employed tube actuators to deploy and retract small fences capable of significantly reducing blade-vortex interaction generated noise on a rotorcraft. Most recently, the variable geometry fan nozzle program has built on the VGC technology to demonstrate improved jet engine performance. Continued maturation of SMA technology is needed in order to develop innovative applications and support their commercialization.

Core Noise Diagnostics of Turbofan Engine Noise UsingCorrelation and Coherence Functions

Cross-correlation and coherence functions are used to look for periodic acoustic components in turbofan engine combustor time histories, to investigate direct and indirect combustion noise source separation based on signal propagation time delays, and to provide information on combustor acoustics. This investigation uses a combustor pressure sensor and a far-field microphone at 130° to study the change in propagation time to the far field of the direct and indirect combustion noise signal using a generalized cross-correlation function. Results are presented as a function of the cutoff frequency of the low-pass filter used to create the generalized cross-correlation function and engine operating condition. The filtering procedure used produces no phase distortion. As the low-pass-filter frequency is decreased, the travel time increases. The indirect combustion noise signal travels more slowly, because the entropy in the combustor moves with the flow, which has a low velocity. The indirect combustion noise signal travels at acoustic velocities after reaching the turbine and being converted into an acoustic signal. The direct combustion noise is always propagating at acoustic velocities. This is in agreement with previous investigations of delay times using a cross-spectrum phase-angle method with unfiltered signals that found indirect combustion noise to be in the 0―200 Hz frequency range and the direct combustion noise to be in the 200-400 Hz frequency range. Similar results obtained using the cross-spectrum phase-angle method with filtered signals are also shown. These results show that the low-pass filtering can be used with the cross-correlation function to separate this type of dependent source and confirm the cross-spectrum results. Although the results are based on a set of static engine tests conducted for one specific dual-spool turbofan engine configuration, they may lead to a better idea about the acoustics in the combustor turbine-tailpipe system and may help develop and validate improved reduced-order physics-based methods for predicting turbofan engine core noise.

A comparative performance study of vehicle, static engine and synthetically aged autocatalysts using cvs, static engine and synthetic gas rig testing methods

The relative performances of on-road, static engine and furnace aged autocatalysts were measured using three established testing methods. Similar trends were noted between vehicle and static engine evaluation data. Perturbed lambda scan data generated on static engine and synthetic gas rig apparatus also showed good correlation. However, on-road aged catalysts were found to be far more severely deactivated than those subjected to laboratory accelerated ageing.

News and Views

First phase static engine test running of the two RB‐199 Mk.104D engines, which power the EAP (Experimental Aircraft Programme) demonstrator, were successfully completed at British Aerospace Military Aircraft Division Warton, Lancashire.

Experimental Study of Flight Effect on Fan Noise : 2nd, Report, Fan Noise Variations due to Flight Effect

Clarification of fan noise source variation due to a flight effect is important for a turbofan noise reduction and a noise estimation. A fan noise spectral extent and a range of fan operation influenced by the flight effect were studied experimentally using a large turbofan engine and a geodesic 4m diameter semi-sphere inflow control device that can simulate the flight effect on fan noise in static engine tests. The fan noise reduction by the flight effect in a fan approach condition was higher than that in a fan take-off condition, because the fan rotor generates BPF tones by itself near the take-off condition. The highest reduced component in fan noise spectra was a fundamental fan blade passing frequency (BPF) tone, whose reduced level was 20 dB at maximum. Higher harmonic BPF tones and a buzz saw noise were also reduced by the flight effect, but a reduction of broadband noise was not evaluated in the whole range of fan operation.

Correction of fan noise for effects of forward flight

When acoustic measurements are made on a static engine test stand, the data must be corrected for the effects of forward flight to predict correctly the noise characteristics of the engine in flight. A ray tracing approach is used here to relate the static test case to the flight case. The assumptions of isentropic irrotational flow into the fan inlet and a cylindrical shear layer at the fan exhaust lead to slightly different methods for correcting inlet noise and exhaust noise. The forward flight correction method generally involves both an angle and an amplitude correction. The amplitude correction factors for inlet and exhaust noise are the same as that for a dipole and can be as much as 6 dB for a flight Mach number of 0·3. The angle correction for the inlet noise differs from that of the exhaust noise, and both differ from the generally used correction to retarded angle.

Farfield radiation of aft turbofan noise

The farfield radiation of turbofan noise from the inlet of an engine has recently received considerable study. Relatively simple approximate expressions have been developed for single mode radiation as a function of cutoff ratio and for multimodal radiation. Convective effects upon radiation which are essential for interpreting static engine tests have been developed. This information has been essential for the design of inlet suppressors. This paper represents an initial effort to provide the same type of radiation expressions for the noise radiation from the aft duct. The results of approximate aft radiation equations are compared to more exact Wiener‐Hopf radiation results. Refraction as well as convective effects in the multiple flow streams must be considered. The possibility that the acoustic power distribution is altered during transmission is also discussed. This has an impact on the design and testing of aft duct suppressors.

Total Pressure Recovery of Flared Fan Nozzles Used as Inlets

A simple expression has been developed which accurately represents the total pressure recovery of a flared exhaust nozzle when it is used as an inlet. Physically such a situation can arise during the reverse-thrust operation of a variable pitch fan jet. The formula is written in terms of the freestream and duct Mach numbers and contains two empirical coefficients. One coefficient is associated with internal flow losses in the nozzle and can be found from a static engine test. The other coefficient is associated with flow losses in the external field of the nozzle and can be approximated from existing cone drag data. The developed expression produces an excellent fit of measured total pressure recoveries; differences between calculated and experimental values were generally less than 0.5%.