A series of studies have been conducted to determine the flow quality in the NASA Lewis Icing Research Tunnel. The primary purpose of these studies was to document airflow characteristics, including flow angularity, in the test section and tunnel loop. A vertically mounted rake was used to survey total and static pressure and two components of flow angle at three axial stations within the test section (test section inlet, test plane, and test section exit; 15 survey stations total). This infonnation will be used to develop methods of improving the aerodynamic and icing characteristics within the test section. The data from surveys made in the tunnel loop were used to determine areas where overall tunnel flow quality and efficiency can be improved. A separate report documents similar flow quality surveys conducted in the diffuser section of the Icing Research Tunnel. 1

A series of studies has been conducted to determine the existing flow quality in the NASA Lewis 8-by 6-Foot Supersonic/ 9-by 15-Foot Low Speed Wind Tunnel. The information gathered from these studies was used to determine the types and designs of flow manipulators which can be installed to improve overall tunnel flow quality and efficiency. Such manipulators include honeycomb flow straighteners, turbulence reduction screens, corner turning vanes, and acoustic treatments.The flow quality studies were conducted at several locations around the tunnel loop. Pressure, flow angularity, temperature, and turbulence measurements were made with both fixed and translating probes. Flow visualization techniques using both video (smoke and tufts) and still photography (oil flow patterns) were also used in these studies. A large portion of the study focused on the flow entering and exiting the seven-stage axial flow compressor. The flow entering both the 8-by 6-ft and the 9-by 15-ft test sections was also examined in detail. Dynamic pressure measurements were made to determine the operating conditions within the compressor.Previous measurements in the tunnel indicated possible flow disturbances originating in the compressor. To examine this possibility the flow both entering and leaving the compressor was surveyed. Flow visualization at the compressor inlet showed that the north side of the compressor is receiving less flow than the south side. However, since velocity and flow angularity surveys at the compressor exit showed that the flow is evenly distributed at the compressor exit, there is no apparent discontinuity feeding through the compressor due to the nonsymmetric inlet flow. Dynamic pressure transducers in the compressor casing showed no indication of rotating stall.Flow visualization and pressure measurements revealed a large separation region on the north side of the compressor exit tailcone when the tunnel is operating at conditions producing Mach numbers greater than 1.6 in the 8-by 6-ft test section. The flow visualization (tufts) showed separation off the entire downstream third of the north side of the tailcone, and reversed flow areas on the cone surface. Velocity surveys downstream of the tailcone also showed a velocity deficit along the north side of the tunnel coinciding with the separation off the tailcone. However, this velocity deficit is greatly reduced by a single turbulence reduction screen at the inlet to the 8-by 6-ft test section bellmouth. Tn f-A ii rf;nnThe NASA Lewis Research Center 8-by 6-Foot Supersonic/9-by 15-Foot Low Speed Wind Tunnel is shown in Fig. 1, together with 11 measurement locations around the tunnel loop. The tunnel is a continuous flow propulsion wind tunnel. The 8-by 6-ft test section has a Mach number range of 0.36 to 2.0. The 9-by 15-ft low speed test section has a Mach number range of 0 to 0.20. Since the end result of the study is to improve the flow quality in the test sections (primarily the 8-by 6-ft test section), the test conditions are referenced to the Mach n...