Engineering topics
Publications and source records attributed to Soderman, P. T..
Steel- or aluminum-skinned baffles absorb sound from air flowing over them. Because there is no bulk filler, muffler does not collect contaminants. If dirt accumulates on skin, it can be washed away without damaging muffler.
Fan silencers with low-pressure drop have been designed and installed in the NASA Ames 7- by 10-Foot Wind Tunnel No. 1. The silencers are composed of an absorbent lining flush with the wind-tunnel walls upstream of the fan and an absorbent splitter in the duct downstream of the fan. An acoustical insertion loss of 4 to 12 dB was measured between 100 and 1000 Hz. High-frequency performance of the silencers was reduced by test-section noise and by corner vanes which turned the sound waves so they followed the duct axis. Sound below 1.2 kHz diffracted around the 305-mm-long vanes and impinged on acoustically absorbent corner walls. Wind-tunnel flow losses are discussed.
Interagency studies conducted during the last several years have indicated the need to improve full-scale testing capabilities. The studies showed that the most effective trade between test capability and facility cost was provided by repowering the existing Ames Research Center 40- by 80-foot wind tunnel to increase the maximum speed from about 100 m/s (200 knots) to about 150 m/s (300 knots) and by adding a new 24- by 37-m (80- by 120-ft) test section powered for about a 50-m/s (100-knot) maximum speed. This paper reviews the design of the facility, a few of its test capabilities, and some of its unique features.
An investigation was made of the test-section noise levels at various wind speeds in the Ames 7- by 10-Foot Wind Tunnel No. 1. No model was in the test section. Results showed that aerodynamic noise from various struts used to monitor flow conditions in the test section dominated the wind-tunnel background noise over much of the frequency spectrum. A tapered microphone stand with a thin trailing edge generated less noise than did a constant-chord strut with a blunt trailing edge. Noise from small holes in the test-section walls was insignificant.
Several methods for aiding acoustic measurements inside wind tunnels are described; these include: placing acoustic material in the test section, using low tip-speed drive fans for reducing background noise, using proper microphone strut design for low wind noise, using directional receivers, and simulating the reverberant noise fields. The instrumentation and techniques were developed for acoustic research in both a small and a large closed-test-section wind tunnel at NASA Ames Research Center, and may be applicable to many different types of aeroacoustic wind tunnels.
An end-fire microphone array that utilizes a digital time delay system has been designed and evaluated for measuring noise in wind tunnels. The directional response of both a four- and eight-element linear array of microphones has enabled substantial rejection of background noise and reverberations in the NASA Ames 40- by 80-foot wind tunnel. In addition, it is estimated that four- and eight-element arrays reject 6 and 9 dB, respectively, of microphone wind noise, as compared with a conventional omnidirectional microphone with nose cone. Array response to two types of jet engine models in the wind tunnel is presented. Comparisons of array response to loudspeakers in the wind tunnel and in free field are made.
This paper presents and interprets a series of extensive studies of the application of leading-edge serrations as a device for reducing the vortex noise radiated from stationary and rotating airfoils in low Reynolds number flow. In these studies, a variety of serrations were attached at selected locations near the leading edge of stationary and rotating airfoils. The noise levels of the airfoils were reduced considerably with the serrations attached. An explanation of the aeroacoustic flow mechanisms involved is given.
A prototype four-element end-fire microphone array was designed and built for evaluation as a directional acoustic receiver for use in large wind tunnels. The microphone signals were digitized, time delayed, summed, and reconverted to analog form in such a way as to create a directional response with the main lobe along the array axis. The measured array directivity agrees with theoretical predictions confirming the circuit design of the electronic control module. The array with 0.15 m (0.5 ft) microphone spacing rejected reverberations and background noise in the Ames 40- by 80-foot wind tunnel by 5 to 12 db for frequencies above 400 Hz.
A method to determine free-field aircraft noise spectra from wind-tunnel measurements has been developed. The crux of the method is the correction for reverberations. Calibrated loud speakers are used to simulate model sound sources in the wind tunnel. Corrections based on the difference between the direct and reverberant field levels are applied to wind-tunnel data for a wide range of aircraft noise sources. To establish the validity of the correction method, two research aircraft - one propeller-driven (YOV-10A) and one turbojet-powered (XV-5B) - were flown in free field and then tested in the wind tunnel. Corrected noise spectra from the two environments agree closely.
Acoustic effects of serrated brass strips attached near the leading edges of two different size rotors were investigated. The two bladed rotors were tested in hover. Rotor rotational speed, blade angle, serration shape, and serration position were varied. The serrations were more effective as noise suppressors at rotor tip speeds less than 135 m/sec (444 ft/sec) than at higher speeds. high frequency noise was reduced but the low frequency rotational noise was little affected. Noise reductions from 4 to 8 db overall sound pressure level and 3 to 17 db in the upper octave bands were achieved on the 1.52 m (5.0 ft) diameter rotor. Noise reductions up to 4 db overall sound pressure level were measured for the 2.59 m (8.5 ft) diameter rotor at some conditions.
The applicability of the NASA Ames 40- by 80-foot wind tunnel for acoustic research on STOL concepts has been investigated. The acoustic characteristics of the wind-tunnel test section have been studied with calibrated acoustic sources. Acoustic characteristics of several large-scale STOL models have been studied in both the free-field and wind-tunnel acoustic environments. The results of these studies indicate that the acoustic characteristics of large-scale STOL models can be measured in the wind tunnel if the test section acoustic environment and model acoustic similitude are taken into consideration. The reverberant field of the test section must be determined with an acoustically similar noise source. A directional microphone, a phased array of microphones, and extrapolation of near-field data to far-field are some of the techniques being explored as possible solutions to the directivity loss in a reverberant field. The model sound pressure levels must be of sufficient magnitude to be distinguishible from the wind-tunnel background noise.
Wind tunnel tests to determine noise data from scale model research aircraft are discussed. Comparisons are made between data obtained in wind tunnels and results of full scale flight tests. The acoustic measurements for the XV-5B V/STOL fan research aircraft are presented.
An investigation was conducted to determine the flow field and aerodynamic effects of leading-edge serrations on a two-dimensional airfoil at a Mach number of 0.13. The model was a NACA 66-012 airfoil section with a 0.76 m (30 in.) chord, 1.02 m (40 in.) span, and floor and end plates. It was mounted in the Ames 7- by 10-Foot Wind Tunnel. Serrated brass strips of various sizes and shapes were attached to the model in the region of the leading edge. Force and moment data, and photographs of tuft patterns and of oil flow patterns are presented. Results indicated that the smaller serrations, when properly placed on the airfoil, created vortices that increased maximum lift and angle of attack for maximum lift. The drag of the airfoil was not increased by these serrations at airfoil angles of attack near zero and was decreased at large angles of attack. Important parameters were serration size, position on the airfoil, and spacing between serrations.
The applicability of the NASA Ames 40- by 80-ft wind tunnel for acoustic research on STOL concepts has been investigated. The acoustic characteristics of the wind tunnel test section has been studied with calibrated acoustic sources. Acoustic characteristics of several large-scale STOL models have been studied both in the free-field and wind tunnel acoustic environments. The results indicate that the acoustic characteristics of large-scale STOL models can be measured in the wind tunnel if the test section acoustic environment and model acoustic similitude are taken into consideration. The reverberant field of the test section must be determined with an acoustically similar noise source. Directional microphone and extrapolation of near-field data to far-field are some of the techniques being explored as possible solutions to the directivity loss in a reverberant field. The model sound pressure levels must be of sufficient magnitude to be discernable from the wind tunnel background noise.
The YOV-10A Research Aircraft was flown to obtain flyover noise data that could be compared to noise data measured in the 40- by 80- foot wind tunnel at NASA Ames Research Center. The flyover noise measurements were made during the early morning hours on runway 32L at Moffett Field, California. A number of passes were made at 50 ft altitude in level flight with an airplane configuration closely matching that tested in the wind tunnel. Two passes were selected as prime and were designated for full data reduction. The YOV-10A was flown over a microphone field geometrically similar to the microphone array set up in the wind tunnel. An acoustic center was chosen as a matching point for the data. Data from the wind tunnel and flyover were reduced and appropiate corrections were applied to compare the data. Results show that wind tunnel and flight test acoustic data agreed closely.
The aerodynamic characteristics of a full scale executive type jet transport aircraft with a T-tail were investigated in a 40 x 80 ft (12.2 by 24.4 meter) wind tunnel (subsonic). Static, longitudinal, and lateral stability, and control characteristics were determined at angles of attack from -2 deg to +42 deg. The aircraft wing had 13 deg of sweep and an aspect ratio of 5.02. The aircraft was tested power off with various wing leading- and trailing-edge high lift devices. The basic configuration was tested with and without such components as engine nacelles, wing tip tanks, and empannage. Hinge-moment data were obtained and downwash angles in the horizontal-tail plane location were calculated. The data were obtained at Reynolds numbers of 4.1 million and 8.7 million based on mean aerodynamic chord. The model had static longitudinal stability through initial stall. Severe tail buffet occurred near the angle of attack for maximum lift. Above initial stall the aircraft had pronounced pitch-up, characteristic of T-tail configurations. A stable trim point was possible at angles of attack between 30 deg and 40 deg (depending on c.g. location and flap setting). Hinge-moment data showed no regions with adverse effects on stick force. Comparisons of wind-tunnel data and flight-test are presented.
Wind tunnel tests of full scale advancing blade concept rotor system at high advance ratio
Trailing-edge flap system with boundary layer control and external jet augmentation for subsonic jet transport