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Airspeed of Icing Research Tunnel Increased 40% by Fan Blade Change
Microphone Holder for Low-Speed Wind Tunnel Improved
Surface Pressures on Wind Tunnel Models Visualized With Pressure-Sensitive Paint
Hypersonic Tunnel Facility Reactivated
Airspeed of Icing Research Tunnel Increased 40% by Fan Blade Change
The NASA Lewis Icing Research Tunnel (IRT) is the world's largest refrigerated icing tunnel and one of only three such facilities in the United States. In this unique facility the icing conditions encountered by today's aircraft are duplicated to test ice effects on actual aircraft components and models of complete airplanes and helicopters as well as the effectiveness of ice detection and protection methods. Although very difficult to simulate in a wind tunnel, natural icing conditions have been successfully achieved in the IRT by carefully controlling the tunnel air temperature (down to 220 °ree;F) and spraying an air-water mixture into the airstream upwind of the test article. An artificial "cloud" is created in which the liquid water content and droplet size are carefully controlled. The resulting icing conditions so closely simulate actual flight conditions that the Federal Aviation Administration (FAA) recognizes IRT testing as an alternative to flight testing for the certification of new icing instrumentation and ice protection systems.
The test section of the IRT is 6 ft high, 9 ft wide, and 20 ft long, a size that makes it unique among refrigerated wind tunnels. Often, however, a significant amount of this large flow area is blocked when the test model is a full-scale component. Blockages of 20% and more are common in IRT tests, relative to typical blockages of less than 5% in conventional wind tunnels. High blockage significantly reduces the maximum airspeed that the tunnel fan can produce, as illustrated in the graph. Here the downward-sloping lines show how rapidly the maximum airspeed capability declines with increasing blockage by the test model.
The area of the graph of greatest interest to the U.S. aeronautics industry and the Department of Defense (the principal IRT customers) is the crosshatched area centered on an airspeed of 300 mph and a model blockage of 20%. As shown by the lower shaded area, before improved fan blades were installed, the airspeeds in the IRT test envelope fell below the area of greatest customer interest by approximately 30 to 130 mph.
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The photograph shows the set of improved fan blades before they were installed in the IRT drive in November 1993. Blade pitch angle was increased by 5°ree; and tip clearances were reduced. The resulting benefits are shown by the bold line in the graph, which defines the current IRT test envelope. At zero blockage maximum airspeed has been raised 40%, from 300 mph to over 420 mph. More importantly, the test envelope now passes through the customers' area of interest. The potential benefits of these fan blade improvements to certification testing are closer correlation between test and expected service, fewer test points required for the same level of result reliability, clearer evidence of meeting/not meeting certification requirements, and greater flight safety. A special-purpose computer program developed at NASA Lewis (ref. 1) was used to model the aerodynamic behavior of the IRT fan drive and guide the redesign of the blades. This software is now available through NASA's Computer Software Management and Information Center (COSMIC).
Use of the IRT is open to commercial and academic organizations as well as government agencies, both civilian and military. Commercial organizations compensate NASA for testing expenses in accordance with an established fee schedule. Testing falls into three general categories:
Reference
Microphone Holder for Low-Speed Wind Tunnel Improved
NASA Lewis has been engaged in ongoing efforts to reduce the level of background noise recorded in the 9- by 15-Foot Low-Speed Wind Tunnel in order to promote it as a premier acoustic facility for future fan testing. Soon all aircraft engines will need to satisfy stringent noise regulations set forth by the Federal Aviation Administration (FAA). The acoustic band in which engine manufacturers will be required to test (due to scaling effects) was discovered to be within the existing background noise levels of the tunnel. The background noise was effectively masking an acoustic band of interest.
Early test results (March 1993) indicated that a major contributor to the background noise levels was generated by the microphone holders (stands). A parametric study of several types of stands followed. Among those tested was a 1-ft-tall machined aluminum aerodynamic foil on loan from the NASA Ames Research Center. (Ames demonstrated that improving the design of in-flow microphone holders could significantly reduce their contribution to the overall background noise levels. This work was accomplished under the direction of Paul Soderman and Christopher Allen.) The test results showed that the NASA Ames microphone stand generated less self-noise than any previously tested microphone stand.
A team was then formed to determine the optimum stand design appropriate for the 9- by 15-Foot Low-Speed Wind Tunnel. The Ames 1-ft-tall contoured airfoil design was used as a baseline. A mold was made of the Ames stand before it was returned to Ames. Considering weight, stiffness, and acoustic interaction requirements, we decided to fabricate the new microphone stands out of a composite material. The new stands were designed and primarily fabricated in-house. The composite shells were made by NASA Lewis' Wood Model Shop.
The microphone stand hardware consists of a fiberglass shell (.22 in. long for the wall-mounted stands) filled with foam. (See photograph on page 36.) The design was optimized by creating an internal duct within the shell and baseplate for routing the microphone cabling, thus eliminating tones produced by extraneous wires. The mounting baseplate was made of aluminum and sealed to the bottom of the fiberglass shell. A boom extension holds the actual microphone probe away from the stand body. The boom is approximately 18 in. long. A 4-ft-tall stand was also designed and fabricated to operate on the tunnel's traversing hardware.
Acoustic calibration testing has proven that the new stands have significantly reduced the measurable background noise levels in the tunnel test section. The stands are durable and easy to use and maneuver within the test section. The fiberglass construction minimized the cost and has shown no evidence of inducing noise due to lack of stiffness. Richard P. Woodward will present a paper at the 33rd AIAA Aerospace Sciences Meeting and Exhibit, January 1995, detailing the background noise level reduction in the 9- by 15-Foot Low-Speed Wind Tunnel. The general conclusion of the acoustic calibration is that the improved tunnel is capable of measuring in-flight noise levels of future low-noise test articles and provides a background aeroacoustic environment comparable to state-of-the-art test facilities.
Lewis contacts: Richard P. Woodward, (216) 433-3923; Bonnie A. Kee-Bowling, (216) 433-5681
Headquarters program office: OA
Surface Pressures on Wind Tunnel Models Visualized With Pressure-Sensitive Paint
Surface pressures on wind tunnel models are one of the most frequently made measurements in aeronautical testing. Until recently, pressure measurements were limited to at most several thousand predetermined locations. The relatively new technique of photoluminescent paint fills in the data gaps that result from using only discrete pressure measurement systems. By using a combination of these two types of systems, accurate pressure mappings can be obtained over entire wind tunnel model surfaces.
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Pressure-sensitive paint (PSP) measurement is a nonintrusive, noncontact technique that creates a continuous pressure map of the surface of a wind tunnel model. To obtain the desired readings, a fluorescent coating of PSP is sprayed on the model. During testing the coating is illuminated with blue light and emits light at a longer wavelength, in this case yellow-orange. The amount of luminescence (or brightness) change depends on the partial pressure of oxygen molecules in air at the model surface. The lower the pressure, the brighter the area or the less quenching done by the oxygen molecules. Areas of higher pressure appear dimmer owing to more quenching. This emitted luminescence is detected by a precision, digital, charge-coupled-device camera, and the acquired images are stored on a personal computer. The measured luminescence images can then be calibrated to give absolute pressure readings. The luminescent/oxygen quenching process is self-refreshing, eliminating the need for cleaning or repainting the surface after each test condition.
The calibration procedure requires the acquisition of two images, a reference and the run or test image. The reference image is taken with a constant pressure on the surface of the test article, typically at atmospheric pressure with no wind hitting the surface. This image is used to normalize or ratio out nonuniformities in the illumination system as well as inconsistencies in the application of the paint. The run image is acquired with the wind tunnel flowing air at the desired test condition. Before a ratio of the two images can be performed, any model movement or deformation needs to be corrected so that the two images overlap each other pixel for pixel. The images are then ratioed, the reference image is divided by the run image, and the PSP calibration is applied to the ratioed image-resulting in the pressure image.
The PSP technique has been used in Lewis wind tunnels with air speeds ranging from 30 knots to Mach 4.0. These tests have shown that pressure changes as low as 0.02 psi can be seen using PSP. The pressure-sensitive paint used at Lewis was developed by McDonnell Douglas Aerospace (MDA) chemists, originally for their internal use. MDA has since provided the PSP to the research community for use in other wind tunnel environments. The PSP system was developed at Lewis from commercially available components, including the digital camera, the personal computer, and the software required for acquisition and data reduction. This technique can now be used by our research customers to complement existing discrete pressure measurement systems.
PSP gives a visual understanding of the pressure phenomena occurring on the surface of aeronautical wind tunnel models, changing the procedures used in wind tunnel test programs. Depending on the results of PSP data taken early in a test program, test matrices and methods can be varied to investigate any anomalies not consistent with pretest predictions.
Lewis contacts: Timothy J. Bencic, (216) 433-5690;
Brian P. Willis, (216) 433-2176
Headquarters program office: OA
Hypersonic Tunnel Facility Reactivated
NASA Lewis' Hypersonic Tunnel Facility, located at Plum Brook Station, in Sandusky, Ohio, has undergone a 4-year reactivation that was completed in May 1994 with an integrated systems test. This reactivation has made available to the nation a powerful hypersonic research tool that provides clean airflow to the test article. Its capability can be expanded to Mach 10 by adding an in-line heater.
From 1971 to 1974 the facility was used to test a model of the hypersonic research engine, an axisymmetric, dual-mode combustion, hydrogen-fueled scramjet engine. Fifty-two tests were successfully completed, for a total test time of 112 min.
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The facility develops its clean air capability by using a 3-MW inductively heated storage heater. Graphite blocks 6 ft in diameter and 2 ft high with 1100 drilled-through holes are heated to 4500 °ree;R by 4600-psig gaseous nitrogen passed through them. Gaseous oxygen is added to create the synthetic air that is expanded through one of the presently available nozzles at Mach 5, 6, or 7. The free-jet, blowdown tunnel has a 42-in.-diameter, 10-ft-long test section. After the air is passed over the test article, it is exhausted to a diffuser section that is pumped by a large steam ejector. The facility controls and instrumentation system were completely updated.
The facility's support systems have not been reactivated. The facility has extensive hydrogen handling capability and can supply hydrogen as a liquid, at ambient temperature, or heated to 1200 °ree;F. It also has a single-pass schlieren system and a model injection system with a single-axis thrust table for thrusts to 8500 lb. Water for cooling the models is available.
The Hypersonic Tunnel Facility stands ready to provide knowledge in many advanced fields, including (but not limited to) hypersonic ramjets and scramjets, combined-cycle engines, advanced rocket engines (with or without air augmentation), direct-coupled combustor tests, and structures thermodynamic testing.
Lewis contact: William D. Pack, (419) 621-3347
Headquarters program office: OA
Last updated 1995
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