This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP avg ) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow. Nomenclature a = distance between the inlet lip highlight station and the inlet throat station (see fig. 3), in. A C = inlet capture (highlight) area; area enclosed by inlet highlight (see fig. 3 for highlight definition) and tunnel wall, in. = distance between the inlet highlight height and the inlet throat height (see fig. 3), in. D 2 = duct diameter at AIP (see fig. 3), in. DPCP avg = average SAE circumferential distortion descriptor DPRP i = SAE radial distortion descriptor for ring i on AIP total-pressure rake h = height of vortex generator, in. H = boundary-layer shape factor, δ * /θ H i = height of inlet throat (see fig. 3), in. H max = maximum height of inlet cowl (see fig. 3), in. ΔH = distance between inlet throat centroid and duct exit centroid (see fig. 3), in. i = ring number on AIP total-pressure rake, value increases from 1 in hub region to 5 in tip region L = length of inlet duct from throat to AIP (see fig. 3 https://ntrs.nasa.gov/search.jsp?R=20060004990 2018-05-12T20:30:27+00:00ZAmerican Institute of Aeronautics and Astronautics 2 P t = free-stream total pressure, psi P t,2,avg = area weighted average total pressure at AIP P t,2,avg /P t = inlet recovery pressure ratio Rn/ft = R...
Boundary layer ingestion (BLI) is explored as a means to improve overall system performance for a Blended Wing Body configuration. The benefits of BLI for vehicle system performance benefit are assessed with a process derived from first principles suitable for highly-integrated propulsion systems. This performance evaluation process provides a framework within which to assess the benefits of an integrated BLI inlet and lays the groundwork for higher-fidelity systems studies. The results of the system study show that BLI provides a significant improvement in vehicle performance if the inlet distortion can be controlled, thus encouraging the pursuit of active flow control (AFC) as a BLIenabling technology.The effectiveness of active flow control in reducing engine inlet distortion was assessed using a 6% scale model of a 30% BLI offset, diffusing inlet. The experiment was conducted in the NASA Langley Basic Aerodynamics Research Tunnel with a model inlet designed specifically for this type of testing. High mass flow pulsing actuators provided the active flow control. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion was determined by 120 total pressure measurements located at the aerodynamic interface plane. The test matrix was limited to a maximum freestream Mach number of 0.15 with scaled mass flows through the inlet for that condition. The data show that the pulsed actuation can reduce distortion from 29% to 4.6% as measured by the circumferential distortion descriptor DC60 using less than 1% of inlet mass flow. Closed loop control of the actuation was also demonstrated using a sidewall surface static pressure as the response sensor.
This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCP avg ) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow. Nomenclature a = distance between the inlet lip highlight station and the inlet throat station (see fig. 3), in. A C = inlet capture (highlight) area; area enclosed by inlet highlight (see fig. 3 for highlight definition) and tunnel wall, in. = distance between the inlet highlight height and the inlet throat height (see fig. 3), in. D 2 = duct diameter at AIP (see fig. 3), in. DPCP avg = average SAE circumferential distortion descriptor DPRP i = SAE radial distortion descriptor for ring i on AIP total-pressure rake h = height of vortex generator, in. H = boundary-layer shape factor, δ * /θ H i = height of inlet throat (see fig. 3), in. H max = maximum height of inlet cowl (see fig. 3), in. ΔH = distance between inlet throat centroid and duct exit centroid (see fig. 3), in. i = ring number on AIP total-pressure rake, value increases from 1 in hub region to 5 in tip region L = length of inlet duct from throat to AIP (see fig. 3 https://ntrs.nasa.gov/search.jsp?R=20060004990 2018-05-11T22:26:25+00:00ZAmerican Institute of Aeronautics and Astronautics 2 P t = free-stream total pressure, psi P t,2,avg = area weighted average total pressure at AIP P t,2,avg /P t = inlet recovery pressure ratio Rn/ft = R...
Progress on an experimental effort to quantify the instability mechanisms associated with roughness-induced transition in a high-speed boundary layer is reported in this paper. To simulate the low-disturbance environment encountered during high-altitude flight, the experimental study was performed in the NASA-Langley Mach 3.5 Supersonic LowDisturbance Tunnel. A flat plate trip sizing study was performed first to identify the roughness height required to force transition. That study, which included transition onset measurements under both quiet and noisy freestream conditions, confirmed the sensitivity of roughness-induced transition to freestream disturbance levels. Surveys of the laminar boundary layer on a 7• half-angle sharp-tipped cone were performed via hot-wire anemometry and pitot-pressure measurements. The measured mean mass-flux and Mach-number profiles agreed very well with computed mean-flow profiles. Finally, surveys of the boundary layer developing downstream of an isolated roughness element on the cone were performed. The measurements revealed an instability in the far wake of the roughness element that grows exponentially and has peak frequencies in the 150 to 250 kHz range.
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