Airframe Noise Test Results from the QTD II Flight Test Program

There is an increasing interest in studying supersonic jet noise due to the immense noise level and poor understanding of the generation mechanisms. In many earlier studies, some nonuniform nozzle geometries were proposed to attain noise reductions in supersonic jets. However, the effectiveness and efficiency of noise reduction vary with the nozzle geometries and the fundamental mechanisms responsible for noise reduction are still unclear due to limitations of experimental measurement techniques for supersonic jet flows. Thus the present effort aims to study the noise radiated from different nonuniform nozzles to provide an experimental database for detailed characterizations. An equally important goal is to study acoustic properties in the near-field region to better understand the effect of nozzle exits on the generation of different noise components. In the present study, six lip-modified nozzles were designed and their detailed acoustic fields are compared to that of the circular baseline nozzle. Two different stepped nozzles were firstly tested at over- and perfectly- expanded conditions. The far-field noise spectrum analyses show that the stepped exits play an insignificant role in altering noise emissions at perfectly-expanded condition. At over-expanded condition, however, the longer stepped nozzle produces significant noise reductions at the sideline and upstream quadrants, while the shorter stepped nozzle does not. To explore the physics responsible for the observed noise reduction, detailed near-field acoustic measurements were conducted to locate the sources of different noise components. The location, extent and intensity of large-scale turbulent noise are slightly modified by stepped nozzles at different azimuthal angles. More importantly, no prominent behaviour of broadband shock noise appears to exist in the near-field of the longer stepped nozzle. Subsequently, cross-correlation technique was employed to study the effects of nozzle geometry on the spatial structure of acoustic field. Baseline and stepped nozzles lead to large coherence noise sources at far downstream regions and the noise coherence region appears to be enlarged by the stepped nozzles, especially for the longer stepped nozzle. An experimental investigation for the stepped nozzles was also conducted at under-expanded conditions. When compared to the circular baseline nozzle, far-field acoustic measurements show that stepped nozzles lead to significant noise reductions at certain polar and azimuthal angles. Detailed spectral analysis shows that the noise reduction is mainly due to reduction in broadband shock associated noise and elimination of jet screech phenomenon. Subsequently, qualitative schlieren technique was employed to visualize the resulting shock structures and the results show that shock cell structures within the jets are asymmetrically modified by stepped nozzles. Quantitative schlieren measurements were performed to correlate the shock structures and density gradient fields with the resulting noise components. The results also show that the reduction in broadband shock associated noise is related to the lower shock strengths, as demonstrated by the density gradient profiles. Additionally, acoustic fields of bevelled, double-bevelled and stepped nozzles were compared using detailed far-field acoustic measurements at different jet expansion conditions. At perfectly-expanded condition, the lip-modified nozzles lead insignificant effect on noise emission, while they generate prominent noise reduction at forward quadrant for imperfectly expanded jets. Particually, the S60 nozzle produces the highest noise reduction of 5.5dB and 6dB at conditions of NPR=2.8 and 4, respectively. At under-expanded condition of NPR=5, the B60 is the optimal nozzle configuration where it generates the maximum noise reduction of 9dB at ϕ=90° and θ=180°. It should be mentioned that the studied nozzles occasionally generate noise tradeoff at shallow polar angles, i.e. ϕ≤50°, where they produce comparable or slightly higher noise level to the baseline nozzle. Finally, the effects of nozzle trailing shape on screech radiation are also investigated through acoustic measurements at an under-expanded condition of NPR=5. Bevelled nozzles eliminate jet screech by producing asymmetric shock structures which results in instability waves within the jet column, as well as mismatches in phase and amplitude. Additionally, double-bevelled nozzles produce symmetric shock structures and amplified screech intensity, even when jet-mixing effects have been significantly enhanced. To study the screech dynamics, wavelet transform was employed to analyze the near-field microphone signals. Jet screech and harmonics of different nozzles are highly unsteady and undergo non-periodic and stochastic temporal variations. Subsequently, near-field acoustic measurements were conducted to locate the screech sources. Double-bevelled nozzles have significant impact upon screech intensity levels and peak noise locations, as well as conferring different changes along different measurement planes.

Noise reduction concepts of lowering the jet Mach number and using the inverted velocity profile at fixed thrust and mass flow have been assessed by large-eddy simulations. This assessment involves a baseline jet that has a relatively high jet Mach number near 1.5, and Low-Mj jets that have jet Mach numbers barely above 1.0. The Low-Mj jets have either a uniform or inverted inlet velocity profile. Both noise reduction concepts can reduce the noise in the downstream direction. The combination of these two methods can achieve a noise reduction up to 3.5dB near the peak radiation direction. In addition, because the Low-Mj jets generate very low shock-associated noise, they produce a substantial noise reduction in the upstream forward) direction. However, the inverted velocity profile elevates the noise level at both the upstream and small aft angles. This noise elevation gets worse as the velocity ratio increases. Lowering the jet Mach number at the fixed thrust and mass flow reduces the turbulent kinetic energy. The inverted velocity profile further reduces the turbulent kinetic energy, but it elevates the high-frequency noise source intensities near the nozzle exit. These two competing effects produce a double peak pattern in the far-field spectral sound pressure levels in the peak noise direction.

The objective of this research is to examine how to design low-airframe-noise transport aircraft using Multidisciplinary Design Optimization (MDO). This involves optimizing aircraft to minimize maximum-take-off-weight, while constraining noise at the approach condition. A design methodology which incorporates noise as a design constraint into an MDO formulation is presented. An MDO framework was designed by integrating aircraft conceptual design tools previously developed at Virginia Tech with the Aircraft Noise Prediction Program (ANOPP). Design studies are presented for cantilever wing and Strut-Braced Wing (SBW) transport aircraft with 300 passengers and a 7,700 nm range. The results show that reducing airframe noise by reducing approach speed alone, will not provide significant noise reduction without a large performance and weight penalty. Therefore, more dramatic changes to the aircraft design are needed to achieve a significant airframe noise reduction. Another study showed that the trailing-edge (TE) flap can be eliminated, as well as all the noise associated with that device, without incurring a significant weight and performance penalty. If noise due to the leading-edge (LE) slats and landing gear are reduced, which is currently being pursued, the elimination of the flap will be very significant as the clean wing noise will be the next ‘noise barrier’. Lastly, an airframe noise analysis showed that a SBW aircraft, with short fuselage-mounted landing gear, could have a similar or potentially a lower airframe noise level than a cantilever wing aircraft.

The landing gear system is the dominant airframe noise source for most modern commercial aircraft during approach. This manuscript reports the results from the ALLEGRA (Advanced Low Noise Landing (Main and Nose) Gear for Regional Aircraft) project. This project assessed the performance of several highly realistic low–noise technologies (LNTs) applied to a detailed full–scale nose landing gear (NLG) model in aeroacoustic wind–tunnel experiments. Four individual low–noise concepts tested, namely a ramp door spoiler, a solid wheel axle fairing, wheel hub caps, and multiple perforated fairings. Combinations and small variations of some of these LNTs were also evaluated. The use of multiple planar microphone arrays allowed for the application of 2D and 3D acoustic imaging algorithms to assess the location and strength of the noise sources within the NLG system in different emission directions for each configuration. The wheel axle, the inner wheel hubs, the steering pinions and the torque link were identified as the noisiest NLG elements. The solid wheel axle fairing was the most effective individual LNT, and it improved its performance when applied in combination with the ramp door spoiler and wheel hub caps, reaching overall noise reductions of more than 4 dBA.

An attempt to evaluate the impact noise due to rail joints was carried out by using the scale model tests. The derivation of law of similarity related to wheel/rail noise was made and actual phenomena can be estimated by the scale model tests. Through the measurement with an impact hammer, the vibratory characteristics of track and wheel are estimated in the scale model test with the law of similarity and they are compared with that of the actual track and wheel. It is found that overall trends are well-estimated. In the running test it is shown that the acoustic power of the impact noise is proportional to the power of 2.3 of wheel speed and it is possible that impact noise due to rail joints is simulated by the scale model test. To investigate a potential for impact noise reduction, the relation between impact noise and track parameters, e.g. a soft rail pad, is examined in the scale model test. It is indicated that soft rail pad leads to the reduction of the impact noise and sleeper vibration, though rail vibration is increased. The scale model test may also evaluate the properties associated with rolling noise. And in the vibration of the rail and sleeper the results obtained in the scale model test show a good agreement with those in the field test.

Flyover noise of modern aircraft in their approach configuration can be dominated by airframe noise. Relevant flow noise sources are high lift devices and landing gears. Therefore aeroacoustic tests were performed in the German Dutch Wind Tunnel employing full-scale landing gears of an A340 aircraft. Farfield noise characteristics were determined for both a nose- and a main landing gear, the latter in a 4-wheel and a 6-wheel mock-up configuration, at wind speeds ranging from 50 to 78 m/s. DNW's 3m acoustic mirror and a planar microphone array served for source localization and ranking. Excess tone noise was detected and originated from pin-cavities. Broadband noise levels were confirmed to increase with the 6th power of flow velocity. Noise directivities feature weak level maxima both in the forward and rear arc. Based on results from earlier testing on a smaller A320 gear the effects on noise of gear size and configuration were evaluated. By covering different gear components with streamlined fairings an overall noise reduction on the order of 3 dB was achieved.

In this study, CFD simulations were conducted to investigate the noise reduction effects of solid- and porous-plates, as well as their combination and the side-edge effect of the porous-plates, when installed on the side braces of a main landing gear. Initially, a noise reduction study was conducted using a simplified infinite-long side brace-only model. It was shown that the distant porous-plate with both side-edges bent forward was highly effective in reducing the noise by reducing pressure fluctuation observed in the wake generated from the flat side-edges. Additionally, the use of porous side-edges bent forward attached directly to the side brace was also shown to be effective in reducing noise, which has the advantage of being easier to install on the side brace. Subsequently, the effectiveness of these two promising noise reduction methods was evaluated through their application to a three-dimensional geometry of the side brace. It was confirmed that both plate devices with the porous side-edges bent forward effectively reduce the noise level.

Control strategies for aircraft airframe noise reduction

Effect of noise reducing components on nose landing gear stability for a mid-size aircraft coupled with vortex shedding and freeplay

D EFINED as the “nonpropulsive noise of an aircraft in flight” in [1], noise produced by airframe components becomes significant during the approach-to-land phase of aircraft operations. Research efforts on identifying and understanding the noise sources associated with the individual airframe components, in particular, high-lift systems, have been undertaken by many researchers [2]. It has been confirmed experimentally in [3–5], and analytically by using empirical methods [6], that both leading-edge slats and trailing-edge flap side edges are the most important noise contributors from the high-lift systems to overall noise levels of an aircraft in approach. The high-lift noise, intuitively thinking, is configuration dependent. Any increase in the deflection angle of slat and flap will lead to an increase in overall noise level. The larger the deflection angles are, the higher the noise level will be. This is due to the fact that the highlift components are more loaded when their deflection angles are increased, generatingwhat is referred to as the loading noise. Further, higher deflection angles may induce stronger vortex shedding and turbulent wake downstream which causes higher noise. Such intuitions were supported by experiments conducted at NASA Langley Research Center in 1998 and 2002 [7], which revealed that slat noise is generated due to its configuration setting angles. Khorrami et al. [8] hypothesized with 2-D unsteady computation that the vortex shedding at the slat trailing edge was responsible for the additional highfrequency spike observed only at higher slat angles. An acoustic whistling mechanism at the slat gap due to vortex shedding was later proposed by Tam and Pastouchenko [9] and Agarwal [10]. The theory supported that the simple vortex shedding and feedback mechanism resonating with gap tone frequency caused noise amplification and propagation. Similar observations were made for flap noise, as reported in [11,12]. Both references reported that the measured noise level at higher flap deflection angles exceeded the predicted noise level, which suggests that there might be an additional source contributor, most likely vortex instabilities, as a result of its interaction with the nearby flap upper surface. It may be noted that such phenomenon was not seen at lower settings of flap angles. These observations may infer that high-lift device noise level is directly affected by the configuration setting angles of slats and flaps. In general, high-lift device noise sources are present at wide frequency ranges of low-to-mid andmid-to-high. At a fixed observer location, under a constant flow speed, slat noise is dominant in the low-to-midfrequency range, and flap side-edge noise is dominant at the mid-to-high-frequency range as shown by both Guo and Joshi [13] and Guo et al. [14]. In this study, the proposed technique uses reduced deployment angles of both high-lift devices simultaneously to achieve the most effective noise reduction in different frequency domains. Aerodynamics performance also has to be carefully monitored to ensure no significant loss is present. The objective of the current Note is to examine the effectiveness of a microtab device in airframe noise reduction. Reducing deployment setting angles of both slats and flaps, the associated noise can be alleviated. Loss in aerodynamic lift, as a result of the reduced highlift settings, is compensated by use of a microtab device. The microtab device can be viewed as a small spanwise strip located at the pressure side of flap near the trailing edge, similar to a Gurney flap (Fig. 1). While deployed normal to local airfoil surface, the microtab device alters the local flowfield which leads to increased effective camber, and hence improves lift-to-drag ratio, similar to that of a Gurney flap introduced by Liebeck [15]. In the present work, a threedimensional numerical analysis is presented and compared to sound pressure level (SPL) of the proposed microtab configuration at a specified far-field location to that of the baseline, that is, conventional configuration in approach, to see if any reduction in noise can be achieved.

\n Over recent years, airframe noise has become one of the most important research topics in the aeronautic field, due to reductions in engine noise with the development of high-bypass technology and the necessity to meet the new standards established by regulation entities. It is caused by fluid-structure interaction and pressure fluctuations caused by gaps and discontinuities. Slat is one of the major airframe noise sources due to the complex flow in the gap between it and the wing main element. Although some noise reduction techniques have been applied to the slat, the relation between the slat characteristics and the noise generation has not been completed understood. This study addresses 21 variations in six different parameters of the slat internal geometry for the establishment of their relation with noise generation. Modifications were focused on the cusp and trailing edge regions and their angle, length and thickness were varied. Comparisons between numerical and experimental results of some configurations validated the simulation methodology. Numerical studies were performed by PowerFLOW 5.3® commercial software and the results included transient analysis, as the far-field spectrum, turbulence in the slat cove, power spectra density contours in both flow around the slat and slat surface, and spectra of the pressure fluctuations on the slat surface, and stationary results as pressure coefficient distribution and lift and drag coefficients of the wing. On the other hand, experimental tests were conducted in a closed-section wind tunnel at the Laboratory of Aerodynamics (LAE) of the São Carlos School of Engineering and consisted in measurements of static pressure in the wall model through a ScaniValve and far-field noise through a phased array antenna of microphones, whose post-processing was completed by convectional and CLEAN SC beamforming methodologies. The good agreement in both aerodynamic and aerocoustic results proved the validity of the numerical set-up and the reliability of numerical results. Among all modified parts, the cusp length exerted the major influence on the slat noise and changed significantly the flow characteristics around the slat. However, the angle and thickness of the cusp showed no effect on the aerodynamic and aerocoustic properties of the wing. Some relationships were observed between the slat noise and characteristics of the fluid around the slat, which were constant among all slat modifications.\n

The main landing gears of large commercial aircraft are an important contributor to the total aircraft noise during landing approach. Within a European financed research project called SILENCER (“Significantly Lower Community Exposure to Aircraft Noise”) a study in “advanced low noise landing gear design” was performed to develop operational landing gears that take into account aeroacoustic constraints early in the design stage. In full scale tests a good level of noise reductions was achieved relative to the equivalent conventional landing gears. For the main landing gear however, there was evidence that further noise reductions might be achievable by refining the design of the bogie. This paper describes a follow- on study at 0.25 model scale to investigate in detail some of the design parameters that affect the noise of the bogie. The test results show that the distribution of sources on the bogie is sensitive to the detailed design, but that achieving reductions in total noise is not straightforward. Analysis of flow patterns indicated by the noise data and flow visualization does however lead to a number of ways of adapting the design of the gear to provide significant noise control.

<div class="htmlview paragraph">Students and faculty of the Mechanical Engineering Department at Texas Tech University have developed two add-on devices that can significantly reduce pickup truck drag without limiting vehicle utility. One device is a wing structure mounted behind the top rear of the cab; the other device is a cover over the rear portion of the pickup bed. Both devices have demonstrated 5% to 6% drag reductions. Further, the rear bed cover reduces vehicle lift by 30%. Most significant, scale model tests have indicated drag reduction on the order of 17% and lift reduction of 30% when the wing and bed cover are used together.</div> <div class="htmlview paragraph">Work to date has involved scale model tests in a low speed wind tunnel and full scale tests in the Lockheed wind tunnel. In addition, computational fluid dynamic studies have been performed to aid in designing the test program and interpreting the test results.</div>

Commercial aircraft today require efficient high-lift and control systems on the wings to reduce the drag in flight or decrease the take-off and landing speeds. Morphing mechanisms are one approach for improved high-lift systems. In most cases the objective function is an increased lift to drag ratio or the noise reduction. On closer examination control systems as well as morphing mechanisms are located in a certain wing segment. The transition between a moving wing part and the fixed wing is a step, which creates additional vortices. This segments the wing in span-wise direction and reduces the efficiency. A flexible skin between a moving and a fixed wing parts smooths the contour and minimize the efficiency reduction of the wing. A full scale demonstrator of a wing segment was manufactured with two flexible skin designs. The first subcomponent connects a morphing leading edge with a rib of the wing over a span of one meter. The skin is a material mix of ethylene-propylene-diene monomer (EPDM) rubber and fiberglass-reinforced plastic. The rubber is the basis of the skin and the glass-fiber is added as local skin stiffeners in the form of strips in chord-wise direction. The second subcomponent blends the aileron with a rib of the wing in a triangular design. The connection of three different hinges realizes a morphing triangle, which is loaded in an in-plane shear only state of stress in each aileron position. The core of the triangle is a 3D printed structure, which is free in shear. The covering skin is a combination of EPDM with carbon fibers oriented in +/−30° direction to obtain shear compliance and to resist the loads on the triangle. The deformation of each concept is identified at the demonstrator. Therefore, an optical measurement system scans the surface in the initial and deflected state. The required deformation precision of the concepts differs due to their design. The contour at the leading edge requires a certain shape over the span. The analysis of the skin buckling is one requirement at the transition triangle during the aileron motion. The experimental results show a smooth transition contour at the leading edge and no buckling effects at the triangle. The results can be used for the validation of simulation models. Furthermore, both skin concepts cover the gap between a moving wing segment and a fixed wing part. The elimination of steps in span-wise direction can improve the aero-acoustic behavior along the wing for future aircraft.

Efficient prediction and evaluation of noise performance are crucial to the design and the optimization of landing gear noise. A systematic method is developed to predict and evaluate landing gear noise in the present study, termed as noise spectrum deep learning model (NSDL). In this algorithm, the encoder and decoder are designed to extract noise features and reconstruct noise data. Specifically, a loss function that takes the identification of both broadband noise and tone noise into account is utilized to guide the training direction of the model, aiming to improve the training efficiency and prediction results of the model. Afterward, the mapping relationship between landing gear experimental parameters and noise features is established by multi-layer perceptron. In this study, the detail of the algorithm is analyzed and discussed based on the results of wind tunnel noise experiment and numerical simulation. The results show that the proposed model can effectively and precisely predict landing gear noise under various conditions, including different flow speeds, angles of attack, number of wheels, and heights of the main strut. For the inflow velocity range of 34–75 m/s, the average error of the overall sound pressure level is restricted below 0.83% (0.6 dB). In case only the angle of attack is changed, the average error is reduced to be less than 0.36% (0.3 dB). The prediction results show that the landing gear noise is mainly broadband noise and tone noise mainly appears in the low frequency and intermediate frequency. With the increase in the inflow speed, the broadband noise increases gradually, and the frequency of tone noise gradually shifts to the high frequency band. Additionally, it is found that, for landing gear with four or six wheels, noise is very sensitive to angles of attack and wheel angles of attack. Consequently, the NSDL method shows significant potential in predicting the sound pressure level of landing gears and is expected to improve the efficiency of evaluation and optimization design for noise reduction of landing gear.